Aluminum alloy products having improved property combinations and method for artificially aging same

ABSTRACT

Aluminum alloy products about 4 inches thick or less that possesses the ability to achieve, when solution heat treated, quenched, and artificially aged, and in parts made from the products, an improved combination of strength, fracture toughness and corrosion resistance, the alloy consisting essentially of: about 6.8 to about 8.5 wt. % Zn, about 1.5 to about 2.00 wt. % Mg, about 1.75 to about 2.3 wt. % Cu; about 0.05 to about 0.3 wt. % Zr, less than about 0.1 wt. % Mn, less than about 0.05 wt. % Cr, the balance Al, incidental elements and impurities and a method for making same. The instantly disclosed alloys are useful in making structural members for commercial airplanes including, but not limited to, upper wing skins and stringers, spar caps, spar webs and ribs of either built-up or integral construction.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.14/255,536, filed Apr. 17, 2014, which is a continuation of U.S. patentapplication Ser. No. 12/152,635, filed May 14, 2008, now U.S. Pat. No.8,840,737, which is a continuation-in-part of U.S. patent applicationSer. No. 11/748,021, filed May 14, 2007, now U.S. Pat. No. 8,673,209,all entitled “ALUMINUM ALLOY PRODUCTS HAVING IMPROVED PROPERTYCOMBINATIONS AND METHOD FOR ARTIFICIALLY AGING SAME”, each of which areincorporated herein by reference in their entirety.

BACKGROUND

1. Field of the Disclosure

The present disclosure relates to aluminum alloys, particularly 7000Series (or 7XXX) aluminum (“Al”) alloys as designated by the AluminumAssociation. More particularly to aluminum alloy products useful inmaking structural members for commercial airplanes that are at most 4inches in thickness.

2. Description of the Related Art

The industry demands on aluminum alloys have become more and morerigorous with each new series of aircraft manufactured by the aerospaceindustry. As the size of new jet aircraft get larger, or as currentjetliner models grow to accommodate heavier payloads and/or longerflight ranges to improve performance and economy, the demand for weightsavings in structural components such as wing components continues toincrease.

A traditional aircraft wing structure is shown in FIG. 1 and includes awing box generally designated by numeral 2. Wing box 2 extends outwardlyfrom the fuselage as the main strength component of the wing and runsgenerally perpendicular to the plane of FIG. 1. In wing box 2, upper andlower wing skins 4 and 6 are spaced by vertical structural members orspars 12 and 20 extending between or bridging upper and lower wingskins. Wing box 2 also includes ribs which extend generally from onespar to the other. These ribs lie parallel to the plane of FIG. 1whereas the wing skins and spars run perpendicular to the FIG. 1 plane.

The upper wing cover is typically comprised of a skin 4 and stiffeningelements or stringers 8. These stiffening elements can be attachedseparately by fastening or made integral with the skin to eliminate theneed for separate stringers and rivets. During flight, the upper wingstructure of a commercial aircraft wing is compressively loaded, callingfor alloys with high compressive strength. This requirement has led tothe development of alloys with increasingly higher compressive strengthwhile still maintaining a nominal level of fracture toughness. The upperwing structural members of today's large aircraft are typically madefrom high strength 7XXX series aluminum alloys such as 7150 (U.S.Reissue Pat. No. 34,008), 7449 (U.S. Pat. No. 5,560,789) or 7055aluminum (U.S. Pat. No. 5,221,377). More recently, U.S. Pat. No.7,097,719 discloses an improved 7055 aluminum alloy.

However, the development of ultra-high capacity aircraft has led to newdesign requirements. Due to a larger and heavier wing and high aircraftgross takeoff weights, these aircraft experience high down-bending loadsduring landing producing high tensile loads in the upper wing structuralmembers. While the tensile strength in the current upper wing alloys ismore than adequate to withstand these down-bending loads, their fracturetoughness becomes a limiting design criterion on the inboard portions ofthe upper cover. This has led to a desire for alloys for the upperstructural members of ultra-large aircraft having very high fracturetoughness more akin to that in lower wing skin alloys such as 2324 (U.S.Pat. No. 4,294,625) even if high strength must be sacrificed to someextent. That is, there has been a shift in the optimum combination ofstrength and toughness needed to maximize weight savings in the upperwing structural members of an ultra-large aircraft to significantlyhigher fracture toughness and lower strength.

New welding technologies such as friction stir welding have also openedmany new possibilities for both design and alloy products for use inwing spar and rib components for weight reduction and/or cost savings.For maximum performance of a spar, the part of the spar which joins tothe upper wing skin would have properties similar to the upper skin, andthe part of the spar which connects to the lower wing skin would haveproperties similar to the lower wing skin. This has led to the use of“built-up” spars, comprising an upper spar cap 14 or 22, a web 18 or 20,and a lower spar cap 16 or 24, joined by fasteners (not shown). This“built-up” design allows optimal alloy products to be used for eachcomponent. However, the installation of the many fasteners requiredincreases assembly cost. The fasteners and fastener holes may also bestructural weak links and parts may have to be thickened which somewhatreduces the performance benefit of using multiple alloys.

One approach used to overcome the assembly cost associated with abuilt-up spar is to machine the entire spar from a thick plate,extrusion or forging of one alloy. Sometimes, this machining operationis known as “hogging out” the part. With this design, the need formaking web-to-upper spar and web-to-lower spar joints is eliminated. Aone piece spar fabricated in this manner is sometimes known as an“integral spar”. An ideal alloy for making integral spars should havethe strength characteristics of an upper wing alloy combined with thefracture toughness and other damage tolerance characteristics of thelower wing alloy. Typically, achieving both properties simultaneously isdifficult and requires a compromise between the property requirementsfor the upper skin and for the lower skin. One disadvantage that anintegral spar must overcome is that the strength and toughnessproperties of a thick product used as the starting stock are typicallyless than those of thinner products typically used in a “built-up” spareven if the integral spar is made of the same alloy and temper. Thus,the compromise in properties and the use of thick products for anintegral spar may result in a weight penalty. One thick product alloywhich reasonably meets the property requirements of both an upper andlower spar cap and retains good properties even in thick productsbecause of its low quench sensitivity, is alloy 7085 described in U.S.Pat. No. 6,972,110. Another disadvantage of integral spars, regardlessof alloy, is the high ratio of buy weight (i.e., material which ispurchased) to fly weight (i.e., weight of material flying on theaircraft) known as the “buy-to-fly.” This at least partly diminishes thecost advantages of an integral spar over a built-up spar achievedthrough reduced assembly cost.

However, new technologies such as friction stir welding make furtherimprovements in both weight and cost a possibility. A multi-componentspar joined by friction stir welding or other advanced welding orjoining methods combines the advantages of a built-up and integral spar.The use of such methods allows the use of use of products of lesserthickness as well as the use of multiple alloys, product forms and/ortempers which are optimized for each spar component. This expands thealloy product/temper options and improves the material buy-to-fly as ina built-up spar, while retaining a significant portion of the assemblycost advantage of an integral spar.

U.S. Pat. No. 5,865,911 describes a 7000 series alloy envisaged for useas lower wing skin structural members and for wing spar members ofultra-high capacity aircraft. This alloy exhibited improvements instrength, toughness, and fatigue resistance in thin plate form relativeto incumbent lower wing alloys such as 2024 and 2324 (U.S. Pat. No.4,294,625). Similar properties in strength and toughness have beenobtained in alloy 7085 (U.S. Pat. No. 6,972,110) in thin plate form asshown in Table 1. Either of these alloys in thin product form would beuseful for structural members of a lower wing cover and for the lowerspar cap and web of a multi-component spar joined by mechanicalfastening or welding. These alloys are also suitable for ribapplications in either a built-up or integral design. However, thestrength levels achievable in these alloys are typically insufficientfor use in upper wing structural members of large commercial aircraft.Higher strength is also beneficial for the upper spar cap, spar web andfor ribs provided adequate toughness is maintained.

TABLE 1 Properties of Miyasato alloy (U.S. Pat. No. 5,865,911) and 7085(U.S. Pat. No. 6,972,110) in thin plate form. Property Dir Miyasato(1)7085 (2) UTS (ksi) L 82.1 82.6 LT 81.4 82.2 TYS (ksi) L 76.2 78.0 LT75.4 77.2 Klc, Kq (ksi√in) L-T 47.5 44.0 RT T-L 40.7 35.9 Klc, Kq(ksi√in) L-T 42.0 40.5 −65 F. T-L na 34.3 Kapp (ksi√in) L-T 120.8 128.7RT T-L 94.3 104.4 Kapp (ksi√in) L-T 115.5 106.8 −65 F. T-L 74.7 79.0 Kc(ksi√in) L-T 172.9 165.7 RT T-L 123.9 129.1 Kc (ksi√in) L-T 166.4 140.1−65 F. T-L 79.8 84.8 (1)U.S. Pat. No. 5,865,911: Rolled plate 1.2 inchesthick, 86 inches wide (2) 7085, U.S. Pat. No. 6,972,110; Rolled plate1.5 inches thick, 102 inches wide

Thus, a need exists for ultra-high capacity aircraft for an alloy thathas significantly higher toughness than current alloys used in upperwing structural members while still maintaining an acceptable level ofstrength. Such an alloy would also be valuable for use in the upper sparcap and spar web of a multi-component spar joined by mechanicalfastening or welding as well as for wing ribs of a built-up or integraldesign. While the needs of ultra-high capacity aircraft and wings havebeen specifically discussed such an alloy may also prove beneficial foruse in fuselage applications and on smaller aircraft both in built-upand integral structures. In addition, non-aerospace parts such as armorfor military vehicles may also be made from the instant alloy.

SUMMARY OF THE DISCLOSURE

New aluminum alloy products particularly well-suited for aerospacestructural components are provided. In one aspect, the new aluminumalloys (sometimes referred to herein as the “instantly disclosed alloy”)include from about 6.80 to about 8.5 wt. % Zn, about 1.5 or 1.55 toabout 2.00 wt. % Mg, about 1.75 to about 2.30 wt. % Cu; about 0.05 toabout 0.3 wt. % Zr, less than about 0.1 wt. % Mn, less than about 0.05wt. % Cr, the balance substantially Al, incidental elements andimpurities. The alloy products are about 4 inches thick or less, andsometimes about 2.5 or 2.0 inches thick or less having significantlyhigher fracture toughness than prior art alloys used for theseapplications while maintaining acceptable levels of strength, andvice-versa.

In one approach, an aluminum alloy product is provided. The aluminumalloy of the product consists essentially of from about 6.80 to about8.5 wt. % Zn, about 1.5 or 1.55 to about 2.00 wt. % Mg, about 1.75 toabout 2.30 wt. % Cu; about 0.05 to about 0.3 wt. % Zr, less than about0.1 wt. % Mn, less than about 0.05 wt. % about Cr, the balance beingaluminum, incidental elements and impurities. The aluminum alloy mayexhibit, when solution heat treated, quenched and artificially aged, andin parts made from the products, an improved combination of strength andfracture toughness. In one embodiment, the alloy comprises low amountsof iron and silicon impurities. In one embodiment, the alloy includesnot more than about 0.15 wt. % Fe and not more than about 0.12 wt. % Siimpurities. In one embodiment, the alloy includes not more than about0.08 wt. % Fe and not more than about 0.06 wt. % Si impurities. In oneembodiment the alloy includes not more than about 0.04 wt. % Fe and notmore than about 0.03 wt. % Si impurities. The aluminum may be in theform of rolled sheets, rolled plates, extrusions or forgings. In someembodiments, the alloy product is less than 2.5 or 2.0 inches thick atits thickest point. In some embodiments, the alloy product is from about2.5 inches to 4 inches thick at its thickest point.

In one approach, the aluminum alloy is in the form of a rolled platehaving a thickness of less than 2.5 inches, such as a thickness of notgreater than 2.00 inches. In one embodiment, the aluminum alloy of theplate comprises 6.8-8.5 wt. % Zn, 1.5-2.0 wt. % Mg, 1.75-2.3 wt. % Cu,and up to 0.25 wt. % of at least one of Zr, Hf, Sc, Mn, and V, and up toabout 89.95 wt. % aluminum. In one embodiment, the aluminum alloycomprises 7.5-8.5 wt. % Zn, 1.9-2.3 wt. % Cu, 1.5-2.0 wt. % Mg, up to0.25 wt. % of at least one of Zr, Hf, Sc, Mn, and V, and up to about89.1 wt. % aluminum. In one embodiment, the aluminum alloy comprises7.8-8.5 wt. % Zn, 1.95-2.25 wt. % Cu, 1.7-2.0 wt. % Mg, up to 0.25 wt. %of at least one of Zr, Hf, Sc, Mn, and V, and up to about 88.55 wt. %aluminum. In one embodiment, the aluminum alloy comprises 7.9-8.2 wt. %Zn, 2.05-2.15 wt. % Cu, 1.75-1.85 wt. % Mg, up to 0.25 wt. % of at leastone of Zr, Hf, Sc, Mn, and V, and up to about 88.3 wt. % aluminum. Inone embodiment, the aluminum alloy comprises 7.4-8.0 wt. % Zn, 1.95-2.25wt. % Cu, 1.7-2.0 wt. % Mg, up to 0.25 wt. % of at least one of Zr, Hf,Sc, Mn, and V, and up to about 88.95 wt. % aluminum. In one embodiment,the aluminum alloy comprises 7.5-7.9 wt. % Zn, 2.05-2.20 wt. % Cu,1.8-1.9 wt. % Mg, up to 0.25 wt. % of at least one of Zr, Hf, Sc, Mn,and V, and up to about 88.65 wt. % aluminum. In various ones of theseembodiments, the aluminum alloy may comprise from 0.05 to about 0.3 wt.% Zr, less than about 0.1 wt. % Mn, and less than about 0.05 wt. % Cr.In any of these embodiments, the aluminum alloy may consist essentiallyof the stated ingredients (aside from aluminum), the balance beingaluminum and incidental elements and impurities. In any of theseembodiments, the alloy product may be less than about 2.5 or 2.0 inchesthick at its thickest point.

In one approach, the aluminum alloy is used in the form of a platehaving a thickness of from 2.5 or 3.0 inches or 2.51 inches to about 3.5inches, 3.75 inches or even 4 inches. In one embodiment, the aluminumalloy of the plate comprises 6.8-8.5 wt. % Zn, 1.5-2.0 wt. % Mg,1.75-2.3 wt. % Cu, and up to 0.25 wt. % of at least one of Zr, Hf, Sc,Mn, and V, and up to about 89.95 wt. % aluminum. In one embodiment, thealuminum alloy comprises 7.4-8.0 wt. % Zn, 1.9-2.3 wt. % Cu, 1.55-2.0wt. % Mg, up to 0.25 wt. % of at least one of Zr, Hf, Sc, Mn, and V, andup to about 89.15 wt. % aluminum. In one embodiment, the aluminum alloycomprises 7.5-7.9 wt. % Zn, 2.05-2.20 wt. % Cu, 1.6-1.75 wt. % Mg, up to0.25 wt. % of at least one of Zr, Hf, Sc, Mn, and V, and up to about88.55 wt. % aluminum. In various ones of these embodiments, the aluminumalloy may comprise from 0.05 to about 0.3 wt. % Zr, less than about 0.1wt. % Mn, and less than about 0.05 wt. % Cr. In any of theseembodiments, the aluminum alloy may consist essentially of the statedingredients (aside from aluminum), the balance being aluminum andincidental elements and impurities.

The alloy product may realize improved strength and toughnessproperties. In one embodiment, the alloy product includes a section notmore than about 2.5 inches or 2.00 inches in thickness and has a minimumtensile yield strength in the longitudinal direction and a plane-strainfracture toughness in the L-T direction at or above and to the right ofline A-A in FIG. 3A or FIG. 3B (e.g., the shaded region). In oneembodiment, the alloy includes a section not more than about 2.5 inchesor 2.00 inches in thickness and having a tensile yield strength and anapparent plane stress fracture toughness in the L-T direction at orabove and to the right of line B-B in FIG. 4 (e.g., the shaded region)when measured in a 16-inch wide center-cracked panel having an initialcrack length (2 ao) of about 4 inches and a thickness of about 0.25inch.

In one embodiment, the alloy product includes a section of from about2.00 or 2.5 inches to 3.0 or 3.125 or 3.25 inches in thickness and has atensile yield strength in the LT (long traverse) direction and aplane-strain fracture toughness in the T-L direction at or above and tothe right of line C-C in FIG. 7 (e.g., the shaded region). In oneembodiment, the alloy product includes a section of from about 2.00 or2.5 inches to 3.0 or 3.125 or 3.25 inches in thickness (e.g., at itsthickest point) and has a tensile yield strength in the ST (shorttraverse) direction and a plane-strain fracture toughness in the S-Ldirection at or above and to the right of line E-E in FIG. 9 (e.g., theshaded region).

In one embodiment, the alloy product includes a section of from about2.75, 3.0, 3.125 or 3.25 inches to about 3.5, 3.75 or 4 inches inthickness (e.g., at its thickest point) and has a minimum tensile yieldstrength in the LT direction and a plane-strain fracture toughness inthe T-L direction at or above and to the right of line D-D in FIG. 8(e.g., the shaded region). In one embodiment, the alloy product includesa section of from about 2.75, 3.0, 3.125 or 3.25 inches to about 3.5,3.75 or 4 inches in thickness and has a minimum tensile yield strengthin the ST direction and a plane-strain fracture toughness in the S-Ldirection at or above and to the right of line F-F in FIG. 10 (e.g., theshaded region).

The alloy product may also realize excellent corrosion resistance. Inone embodiment, the alloy product of has an EXCO corrosion resistancerating of “EB” or better. In one embodiment, the alloy productconsistently passes alternate immersion stress corrosion crackingresistance tests at a stress level of 35 ksi for a T74 temper, at astress level of 25 ksi for a T76 temper, and at a stress level of 15 ksifor a T79 temper. In one embodiment, the alloy product consistentlypasses seacoast environment stress corrosion cracking resistance testsat a stress level of 35 ksi for a T74 temper, at a stress level of 25ksi for a T76 temper, and at a stress level of 15 ksi for a T79 temper.In one embodiment, the alloy product consistently achieves an EXCOcorrosion resistance rating of “EB” or better, and consistently passesboth alternate immersion stress corrosion cracking resistance and aseacoast environment stress corrosion cracking resistance tests at astress level of 35 ksi for a T74 temper, at a stress level of 25 ksi fora T76 temper, and at a stress level of 15 ksi for a T79 temper. In oneembodiment, the alloy product consistently achieves an EXCO corrosionresistance rating of “EB” or better, and consistently passes bothalternate immersion stress corrosion cracking resistance and seacoastenvironment stress corrosion cracking resistance tests at a stress levelof 35 ksi for a T74 temper, at a stress level of 25 ksi for a T76temper, and at a stress level of 15 ksi for a T79 temper, and achievesthe above-described tensile yield strength and fracture toughnessproperties. The alloy product may pass other stress corrosion crackingresistance tests as well.

The alloy product may be utilized in a variety of applications. In oneembodiment, the alloy product is an aerospace structural component. Theaircraft structural component may be any of an upper wing panel (skin),an upper wing stringer, an upper wing cover with integral stringers, aspar cap, a spar web, a rib, rib feet or a rib web, stiffening elementsand combinations thereof. In one embodiment, the alloy product is afuselage component (e.g., a fuselage skin) In one embodiment, the alloyproduct is an armor component (e.g., of a motorized vehicle). In oneembodiment, the alloy product is used in the oil and gas industry (e.g.,as pipes, structural components).

The alloy products may be produced by a variety of methods. For example,the component may be made from an alloy product that is welded by fusionor solid state methods to one or more aluminum alloy products made ofsubstantially the same alloy of the same or different temper to make thecomponent. In one embodiment, the alloy product is joined to one or morealuminum alloy products of different alloy composition to make amulti-alloy component. In one embodiment, the product is joined bymechanical fastening. In one embodiment, the alloy product is joined byfusion or solid state welding methods. In one embodiment, the alloyproduct is age formed either alone or after joining to other alloyproducts in the process of making a component. In one embodiment, thealloy product is reinforced by fiber metal laminates or otherreinforcing materials.

Methods of producing aluminum alloys and aluminum alloy products arealso provided. In one approach, a method includes the steps of formingor shaping an aluminum alloy into an aircraft structural component. Themethod may include producing or providing an aluminum alloy, such as analuminum alloy having any of the aforementioned compositions,homogenizing and hot working the alloy by one or more methods selectedfrom the group consisting of rolling, extruding and forging, solutionheat treating the alloy, quenching the alloy, and stress relieving thealloy. The structural component in an artificially aged condition mayexhibit an improved combination of strength and fracture toughness. Inone embodiment, the alloy is less than about 4 inches thick whenquenched. In one embodiment, the method includes age forming thecomponent either alone or after joining to other components.

In one embodiment, the forming or shaping of the structural componentstep includes machining. In one embodiment, the machining is performedafter artificially aging or between one of the aging stages. In oneembodiment, the machining is performed prior to solution heat treatment.

In one embodiment, the shaping or forming of the structural componentstep includes age forming either before or after joining to othercomponents. In one embodiment, at least some of the forming or shapingof the structural component step is performed before or during at leastsome of the artificial aging.

In one embodiment, the alloy is artificially aged by a method comprising(i) a first aging stage within about 150 to about 275° F., and (ii) asecond aging stage within about 290 to about 335° F. In one embodiment,the first aging stage (i) proceeds within about 200 to about 260° F. Inone embodiment, the first aging stage (i) proceeds for about 2 to about18 hours. In one embodiment, the second aging stage proceeds for about 4to about 30 hours within about 290 to about 325° F. In one embodiment,the second aging stage (ii) proceeds for about 6 to about 30 hourswithin about 290 to about 315° F. In one embodiment, the second agingstage (ii) proceeds for about 7 to about 26 hours within about 300 toabout 325° F. In one embodiment, one or both of the aging stagesincludes an integration of multiple temperature aging effects. In oneembodiment, one or both of the aging stages is interrupted in order toweld the part to another component of the same or a different alloy ortemper.

In another embodiment, the alloy is artificially aged by a methodcomprising (i) a first aging stage within about 290 to about 335° F.,and (ii) a second aging stage within about 200 to about 275° F. In oneembodiment, the first aging stage (i) proceeds for about 4 to about 30hours within about 290 to about 325° F. In one embodiment, the firstaging stage (ii) proceeds for about 6 to about 30 hours within about 290to about 315° F. In one embodiment, the first aging stage (i) proceedsfor about 7 to about 26 hours within about 300 to about 325° F. In oneembodiment, one or both of the aging stages includes an integration ofmultiple temperature aging effects. In one embodiment, one or both ofthe aging stages is interrupted in order to weld the part to anothercomponent of the same or a different alloy or temper.

In another embodiment, the alloy is artificially aged by a methodcomprising (i) a first aging stage within about 150 to about 275° F.,(ii) a second aging stage within about 290 to about 335° F., and (iii) athird aging stage within about 200 to about 275° F. In one embodiment,the first aging stage (i) proceeds within about 200 to about 260° F. Inone embodiment, the first aging stage (i) proceeds for about 2 to about18 hours. In one embodiment, the second aging stage (ii) proceeds forabout 4 to about 30 hours within about 290 to about 325° F. In oneembodiment, the second aging stage (ii) proceeds for about 6 to about 30hours within about 290 to about 315° F. In one embodiment, the secondaging stage (ii) proceeds for about 7 to about 26 hours within about 300to about 325° F. In one embodiment, the third aging stage (iii) proceedsfor at least about 2 hours within about 230 to about 260° F. In oneembodiment, the third aging stage (iii) proceeds for about 18 hours ormore within about 240 to about 255° F. In one embodiment, one, two orall of the aging stages includes an integration of multiple temperatureaging effects. In one embodiment, one, two or all of the aging stages isinterrupted in order to weld the part to another component of the sameor a different alloy or temper.

The method(s) may include joining alloy components. In one embodiment,one or more of the components are joined by mechanical fastening. In oneembodiment, one or more of the components are joined by welding. In oneembodiment, the components are welded by electron beam welding. In oneembodiment, the components are welded by friction stir welding. In oneembodiment, a component is fastened or welded to another aluminumproduct to make a multi-alloy and/or multi-temper component.

As may be appreciated, various ones of the above-noted aspects,approaches and/or embodiments may be combined to yield various usefulaluminum alloy products and components. These and other aspects,advantages, and novel features of the disclosure are set forth in partin the description that follows and will become apparent to thoseskilled in the art upon examination of the following description andfigures, or may be learned by practicing the disclosure.

BRIEF DESCRIPTION OF THE DRAWINGS

For a fuller understanding of the instant disclosure, reference is madeto the following description taken in connection with the accompanyingdrawing(s), in which:

FIG. 1 is a transverse cross-sectional view of a typical wing boxconstruction of an aircraft wing;

FIGS. 2A and 2B illustrate embodiments of the instant alloy compositionin terms of the major alloying elements Cu and Zn and Mg and Zn and ascompared to compositions of 7085 and 7055 and 7449 alloy families,respectively;

FIGS. 2C-1, 2C-2, 2D-1, and 2D-2 illustrate various embodiments of thealloy composition of the present disclosure, such as compositions usefulfor producing aluminum alloy plates having a thickness of not greaterthan 2 or 2.5 inches;

FIGS. 2E and 2F illustrate various embodiments of the alloy compositionof the present disclosure, such as compositions useful for producingaluminum alloy plates having a thickness of at least about 2 or 2.5inches;

FIG. 3A is a graph illustrating typical L-T plane strain fracturetoughness K_(Ic) versus minimum longitudinal tensile yield strengths of(i) example alloys A-D in plate form and of a T79 temper, and (ii)several other conventional alloys in thin plate form;

FIG. 3B is a graph illustrating typical L-T plane strain fracturetoughness K_(Ic) versus minimum longitudinal tensile yield strengths of(i) example alloys A-D in plate form and of a T79 temper, and (ii)several other conventional alloys in plate form;

FIG. 4 is a graph illustrating typical L-T plane stress fracturetoughness K_(app) versus actual or measured tensile yield strengths of(i) example alloys A-D in plate form and of a T79 temper and (ii)several other conventional alloys in plate form;

FIG. 5 is a graph comparing the percentage retained strength aftercorrosion exposure in the LT direction of two of the example alloycompositions for three 3rd step aging times of 0, 6 and 12 hours;

FIG. 6 is a graph comparing the percentage retained strength aftercorrosion exposure in the LT direction of an example alloy and a priorart 7055 alloy for a 12-hour 2nd step aging time.

FIG. 7 is a graph illustrating typical T-L plane strain fracturetoughness K_(Ic) versus typical LT tensile yield strengths of plates of(i) example alloy E (having a thickness of 3.125 inches) and of a T74temper, and (ii) several other conventional alloys (having a thicknessof about 3 inches);

FIG. 8 is a graph illustrating typical T-L plane strain fracturetoughness K_(Ic) versus typical LT tensile yield strengths of plates of(i) example alloy F (having a thickness of 4.0 inches) and of a T74temper, and (ii) several other conventional alloys (having a thicknessof about 4 inches);

FIG. 9 is a graph illustrating typical S-L plane strain fracturetoughness K_(Ic) versus typical ST tensile yield strengths of plates of(i) example alloy E (having a thickness of 3.125 inches) and of a T74temper, and (ii) several other conventional alloys (having a thicknessof about 3 inches); and

FIG. 10 is a graph illustrating typical S-L plane strain fracturetoughness K_(Ic) versus typical ST tensile yield strengths of plates of(i) example alloy F (having a thickness of 4.0 inches) and of a T74temper, and (ii) several other conventional alloys (having a thicknessof about 4 inches).

Like reference characters denote like elements throughout the drawings.

DETAILED DESCRIPTION

FIG. 1 is a schematic illustrating a transverse cross-sectional view oftypical wing box construction 2 comprising upper wing skin 4 andstringers 8, lower wing skin 6 and stringers 10, spaced by spars 12 and20. Stringers 4 and 10 can be attached separately by fastening or madeintegral with the skin to eliminate the need for separate stringers andrivets. Typically, two, three or four wing panels 4 or 6 are needed tocover each of the wing upper and lower surfaces depending on aircraftsize and wing design. Even more panels may be required for an integralskin and stringer design. The multiple panels comprising the upper andlower skin are typically joined by mechanical fastening. These jointsadd weight to the aircraft.

The spars may be of a “built-up” design comprised of upper spar cap 14or 22, lower spar cap 16 or 24 and web 18 or 26 joined by mechanicalfastening or they may be of integral one-piece design, each type ofdesign having its own advantages and disadvantages. A built-up sparallows for optimal alloy products to be used for each of the sparcomponents and has improved “buy-to-fly” compared to an integral spar.Typically, the upper spar cap requires high compressive strength whilethe lower spar caps requires less strength but higher damage toleranceproperties such as fracture toughness and fatigue crack growthresistance. An integral spar has much lower assembly costs but itsperformance may be less than for a built-up design since its propertiesare necessarily a compromise between the requirements for the upper skinand lower skin. Also, strength and toughness of a thick product used asa starting stock for an integral spar are typically less those ofthinner products used for a built-up spar.

The wing box also includes ribs (not shown) which extend generally fromone spar to the other. These ribs lie parallel to the plane of FIG. 1whereas the wing skins and spars run perpendicular to said FIG. 1 plane.Like spars, the ribs can also be of a built-up or integral design witheach type having similar advantages or disadvantages as in spars.However, the optimum properties in ribs differ somewhat with highstrength being advantageous for rib feet which connect to the upper andlower wing skin and stringers and higher stiffness being advantageousfor the web of the rib. More typically, wing ribs are of an integraldesign with a compromise in properties between the requirements for therib feet and rib web.

New welding technologies such as friction stir welding and electron beamwelding allows for new structural concepts retaining the advantages ofcurrent built-up and integral designs while minimizing theirdisadvantages. For example, the different wing panels 4 used to make theupper skin can be joined by friction stir welding instead of amechanically fastened joint thereby reducing the weight of the upperskin. Spars and ribs can be made from multiple alloys, tempers and/orproducts optimized for each spar or rib component joined by frictionstir welding, thereby retaining the performance advantage and betterbuy-to-fly of thinner products as in a built-up spar while reducingassembly costs like an integral spar or rib. For example, upper sparcaps 14 and 22 could be made from a high strength alloy or temperextrusion, lower spar caps 16 and 24 from a lower strength damagetolerance alloy or temper extrusion, and the spar webs 18 and 26 from amoderate strength alloy or temper plate, the three components joined byfriction stir welding or electron beam welding. Designs containing amixture of integral and built-up design could be utilized to improve thefail safety and damage tolerance of a component while reducing assemblycost. For example, the upper spar caps 14 and 22 could be joined byfriction stir welding to the spar webs 12 and 20 to reduce assemblycosts while the lower spar caps 16 and 24 could be mechanically fastenedto improve damage tolerance. Further improvements in damage tolerance inbuilt-up, integral welded and structures containing a mixture of bothcould be achieved by reinforcement with fiber metal laminates and otherreinforcing materials as described in U.S. Pat. No. 6,595,467.

The alloy described in U.S. Pat. No. 6,972,110, which has the commercialdesignation 7085, is primarily directed at thicker gauges, generallyfrom 4 to 8 inches or greater where low quench sensitivity is important.Low quench sensitivity is achieved by providing a carefully controlledcomposition which permits quenching thicker gauges while still achievingsuperior combinations of high strength and toughness and corrosionresistance compared to previous thick product alloys such as 7050, 7010and 7040. The careful composition registered as AA7085 includes low Cu(about 1.3 to about 1.9 wt. %) and low Mg levels (about 1.3 to about1.68 wt. %), which are among the leanest levels used for commercialaerospace alloys. The Zn levels (about 7 to about 9.5 wt. %) at whichthe properties were most optimized corresponded to levels much higherthan those specified for 7050, 7010 and 7040. This was against pastteachings that higher Zn content increases quench sensitivity. On thecontrary, the higher Zn levels in 7085 were actually proven to bebeneficial against the slow quench conditions of thick sectioned pieces.U.S. Pat. No. 6,972,110 teaches that a good portion of the improvedstrength and toughness for thick sections of its instant alloy are dueto the specific combination of alloying ingredients.

U.S. Pat. No. 5,221,377 pertains to the 7055 alloy, which is typicallyused for plate and extrusions 2 inches thick or less, and teaches thatreducing Mg levels results in improved fracture toughness. It is alsowidely appreciated in the prior art that increasing strength throughincreased solute content typically results in a reduction in toughness.

The instant alloy is primarily directed at thinner alloy products, about4 inches thick or less, and sometimes about 2.0 or 2.5 inches thick orless, for upper wing structural members of large commercial aircraftincluding wing skins, wing stringers and upper spar caps. Theseapplications will benefit from and in many cases would require higherstrength than can be achieved by the 7085 composition. Likewise, higherstrength may be beneficial in other applications such as spar webs, ribsand other aerospace components. In order to increase strength, the Mgrange of the instantly disclosed alloys is increased to about 1.5 or1.55 to about 2.0 wt. % and the Cu range from about 1.75 to about 2.30wt. %. The Zn range is lowered somewhat to about 6.8 to about 8.5 wt. %.FIGS. 2A and 2B illustrate embodiments of the instant alloy compositionin terms of the major alloying elements Cu and Zn and Mg and Zn and ascompared to compositions of 7085 (U.S. Pat. No. 6,972,110) and 7055(U.S. Pat. No. 5,221,377) and 7449. Suitable compositions of theinstantly disclosed alloy are designated by a rectangular box with solidlines. The compositions of example alloys A-F, described below, are alsoincluded in FIGS. 2A and 2B.

In one approach, the instantly disclosed alloys are in the form of aplate having a thickness of less than 2.5 inches, such as a thickness ofnot greater than 2.00 inches. In one embodiment, the aluminum alloy ofthe plate comprises 6.8-8.5 wt. % Zn, 1.5-2.0 wt. % Mg, 1.75-2.3 wt. %Cu, and up to 0.25 wt. % of at least one of Zr, Hf, Sc, Mn, and V, andup to about 89.95 wt. % aluminum (e.g., as illustrated in FIGS. 2A and2B). In other embodiments, and with reference to FIGS. 2C-1, 2C-2, 2D-1,and 2D-2, the aluminum alloy comprises 7.5-8.5 wt. % Zn, 1.9-2.3 wt. %Cu, 1.5-2.0 wt. % Mg, up to 0.25 wt. % of at least one of Zr, Hf, Sc,Mn, and V, and up to about 89.1 wt. % aluminum (as provided byembodiment 1 of FIGS. 2C-1 and 2C-2). In another embodiment, thealuminum alloy comprises 7.8-8.5 wt. % Zn, 1.95-2.25 wt. % Cu, 1.7-2.0wt. % Mg, up to 0.25 wt. % of at least one of Zr, Hf, Sc, Mn, and V, andup to about 88.55 wt. % aluminum (as provided by embodiment 2 of FIGS.2C-1 and 2C-2). In one embodiment, the aluminum alloy comprises 7.9-8.2wt. % Zn, 2.05-2.15 wt. % Cu, 1.75-1.85 wt. % Mg, up to 0.25 wt. % of atleast one of Zr, Hf, Sc, Mn, and V, and up to about 88.3 wt. % aluminum(as provided by embodiment 3 of FIGS. 2C-1 and 2C-2. In one embodiment,the aluminum alloy comprises 7.4-8.0 wt. % Zn, 1.95-2.25 wt. % Cu,1.7-2.0 wt. % Mg, up to 0.25 wt. % of at least one of Zr, Hf, Sc, Mn,and V, and up to about 88.95 wt. % aluminum (as provided by embodiment 4of FIGS. 2D-1 and 2D-2). In one embodiment, the aluminum alloy comprises7.5-7.9 wt. % Zn, 2.05-2.20 wt. % Cu, 1.8-1.9 wt. % Mg, up to 0.25 wt. %of at least one of Zr, Hf, Sc, Mn, and V, and up to about 88.65 wt. %aluminum (as provided by embodiment 5 of FIGS. 2D-1 and 2D-2). Invarious ones of these embodiments, the aluminum alloy may comprise from0.05 to about 0.3 wt. % Zr, less than about 0.1 wt. % Mn, less thanabout 0.05 wt. % Cr. In any of these embodiments, the aluminum alloy mayconsist essentially of the stated ingredients (aside from aluminum), thebalance being aluminum and incidental elements and impurities.

In another approach, the aluminum alloy is used in a plate having athickness of from about 2.01 inches or 2.51 inches to about 3.5 inches,3.75 inches or even 4 inches. In one embodiment, the aluminum alloy ofthe plate comprises 6.8-8.5 wt. % Zn, 1.5-2.0 wt. % Mg, 1.75-2.3 wt. %Cu, and up to 0.25 wt. % of at least one of Zr, Hf, Sc, Mn, and V, andup to about 89.95 wt. % aluminum (e.g., as illustrated in FIGS. 2A and2B). In other embodiments, and with reference to FIGS. 2E and 2F, thealuminum alloy comprises 7.4-8.0 wt. % Zn, 1.9-2.3 wt. % Cu, 1.55-2.0wt. % Mg, up to 0.25 wt. % of at least one of Zr, Hf, Sc, Mn, and V, andup to about 89.15 wt. % aluminum (as provided by embodiment 1 of FIGS.2E and 2F). In one embodiment, the aluminum alloy comprises 7.5-7.9 wt.% Zn, 2.05-2.20 wt. % Cu, 1.6-1.75 wt. % Mg, up to 0.25 wt. % of atleast one of Zr, Hf, Sc, Mn, and V, and up to about 88.55 wt. % aluminum(as provided by embodiment 2 of FIGS. 2E and 2F). In various ones ofthese embodiments, the aluminum alloy may comprise from 0.05 to about0.3 wt. % Zr, less than about 0.1 wt. % Mn, less than about 0.05 wt. %Cr. In any of these embodiments, the aluminum alloy may consistessentially of the stated ingredients (aside from aluminum), the balancebeing aluminum and incidental elements and impurities.

From the teachings of U.S. Pat. No. 6,972,110, the composition changesin the instantly disclosed alloys would increase the quench sensitivityof the alloy somewhat in comparison to alloy 7085 and this is quitepossibly the case. However, the instantly disclosed alloys likelyretains some of the benefit of the 7085 composition, and in any case,quench sensitivity is less of a concern in the thinner alloy products atwhich the instantly disclosed alloys are directed. The changes incomposition were also expected to have a detrimental influence onfracture toughness both because of the resulting increase in strengthand the higher Mg content. With the Mg range between that of 7085 andexisting upper wing alloys 7055 and 7449, it was believed that thestrength and toughness of the instantly disclosed alloys would fallbetween these alloys. This was indeed the case for strength. However,the combination of strength and fracture toughness of the instantlydisclosed alloys were improved not over only 7055 and 7449 as expected,but, quite surprisingly it was also improved over the 7085 alloy. Thus,the instantly disclosed alloys identify an unexpected “sweet”composition region that offers higher combinations of strength andfracture toughness than exhibited by the incumbent alloys.

The alloy products of the present disclosure can be prepared by more orless conventional practices including melting and direct chill (DC)casting into ingot form and exhibit internal structure featurescharacteristic of ingot derivation. Conventional grain refiners such asthose containing titanium and boron, or titanium and carbon, may also beused as is well-known in the art. Once an ingot has been cast from thiscomposition, it is scalped (if needed) and homogenized by heating to oneor more temperatures between about 800° and about 900° F., or betweenabout 850° to about 900° F. After homogenization, these ingots areworked by, for example, rolling into plate or sheet or extruding orforging into special shaped sections. For most aerospace applications,alloy products made from the instantly disclosed composition have across-sectional thickness of about 4, 3.75 or 3.5 inches thick or less,and sometimes about 2.5 or 2.0 inches thick or less. The product, ifdesired, should then be solution heat treated by heating to one or moretemperatures between about 850° and about 900° F. to take substantialportions, sometimes all or substantially all, of the soluble zinc,magnesium and copper into solution, it being understood that withphysical processes that are not always perfect, probably every lastvestige of these main alloying ingredients will not be dissolved duringsolution heat treatment. After heating to elevated temperatures, asdescribed, the product should be rapidly cooled or quenched to completethe solution heat treatment procedure. Such cooling is typicallyaccomplished by immersion in a suitably sized tank of cold water or bywater sprays. Air chilling may also be used as a supplementary orsubstitute cooling means. After quenching, certain products may need tobe mechanically stress relieved such as by stretching and/or compressionup to about 8%, for example from about 1% to about 3%.

A solution heat treated and quenched product, with or without coldworking, is then considered to be in a precipitation-hardenablecondition, or ready for artificial aging. The practice may be two-stepor three-step practice and for some applications even a single steppractice may suffice. However, clear lines of demarcation may not existbetween each step or phase. It is generally known that ramping up and/ordown from given (or target) treatment temperatures, in itself, canproduce precipitation (aging) effects which can, and often need to be,taken into account by integrating such ramping conditions, and theirprecipitation hardening effects, into the total aging treatment program.Such integration was described in greater detail in U.S. Pat. No.3,645,804, the disclosure of which is fully incorporated by referenceherein.

U.S. Pat. No. 6,972,110, the disclosure of which is fully incorporatedby reference herein, describes a three (3)-step aging practice for the7085 alloy. A 3-step aging practice with the same or similar temperatureranges to that disclosed in the '110 patent may also be used with theinstantly disclosed alloy, but a 2-step practice is also suitable forsome of the principal applications envisioned. The 2-step practice canbe either the low temperature step followed by the high temperaturestep, or vice-versa. For example, a 2-step practice is often utilizedfor upper wing skins and stringers. These components are often ageformed by the aircraft manufacturer to obtain the contour of the wing.During age forming, the part is constrained in a die at an elevatedtemperature usually between about 250 and about 400° F. for several totens of hours, and the desired contour are accomplished through creepand stress relaxation processes. The age forming is often accomplishedin conjunction with the artificial aging treatment, especially duringthe high temperature step at which creep occurs most rapidly. The ageforming is typically done in an autoclave furnace. The autoclave anddies required to age form an aircraft wing panel for a large commercialaircraft are large and expensive and as a result few are employed in themanufacturing process. Thus, it is desirable that the age forming cyclebe as short as practicable while still achieving the required contourand properties in the alloy product so that production throughput ismaximized. A shortening of the third step or its complete elimination isbeneficial in achieving this goal. In a low-high 2-step practice thefirst step can be applied by the alloy producer, further minimizing thetime expended in the age forming process.

The results of SCC studies on the example alloys indicate that the thirdstep can indeed be shortened and even eliminated while meeting the SCCrequirements for upper wing skin and stringers. The 3-step practice for7085 alloys in thick product applications is generally unnecessary forthe instantly disclosed alloys in upper wing and other high strengthapplications for several reasons. For instance, the SCC requirements forupper wing components are less stringent than those for a thick productapplication such as a rib or spar. The upper wing components arepredominantly subjected to compressive stresses while the spar, inparticular the lower portion, is subjected to tensile stresses. Onlytensile stresses contribute to SCC. Also, an integral spar or ribmachined from a thick product can have significant design stresses inthe ST direction. For example, the spar caps of an integral spar madefrom plate are in the L-ST plane of the parent plate. In comparison, theprincipal design stresses in the upper skin and stringer arepredominantly in the L-LT plane, which is less prone to SCC. As a resultof these differences, the minimum SCC requirement in the ST directionfor the incumbent upper wing alloys 7055 and 7449 is 15 or 16 ksiallowing these alloys to be used in the high strength −T79 temper whilethick products for spars, ribs and other applications are typically usedin the lower strength −T76 and −T74 tempers which typically have SCCminimums of 25 ksi and 35 ksi, respectively.

The instantly disclosed alloys are also envisaged for use in amulti-component, multi-alloy spar or rib joined by mechanical fasteningor welding. As already described, these applications will likely havehigher SCC requirements than for upper wing skin and stringers. However,in a multi-component spar made up of thinner products, the grainstructure can be more favorably oriented for SCC resistance than for anintegral spar machined from thick plate. The spar caps, for example, canbe machined from the more SCC resistant L-LT plane of a parent plate orextrusion instead of the L-ST plane. The minimum SCC performance in theL and LT directions is typically greater than 40 ksi, even in the lessSCC resistant high strength tempers, compared to 25 ksi or 35 ksi in theST direction for the lower strength, higher SCC resistant tempers. Thus,it may be the case that the 3rd step aging practice often utilized for7085 alloys can also be shortened or eliminated for the instantlydisclosed alloys even for spar, rib and other applications having moredemanding SCC requirements. Shortening or elimination of the third stepdoes result in a small strength reduction, typically about 1 to about 2ksi. However, it may be the case that this strength reduction can becompensated for by the use of higher strength tempers not practicable inthick products. Even so, for some built-up, integral or multi-componentapplications of the present disclosure, lower strength tempers such asthe −T74 or −T73 may be desirable, either for the additional corrosionresistance provided or for additional improvements in fracturetoughness.

In the case of multi-alloy spar or rib joined by welding, theflexibility in the aging practice exhibited by the instantly disclosedalloys is a desirable characteristic. The welding, either by fusionwelding methods or solid state methods such as friction stir welding,may be performed in an intermediate temper instead of in the final alloytemper as post weld aging is typically desirable to improve the strengthand corrosion properties of the weld. For example, the welding of theinstantly disclosed alloys to another alloy having strength and damagetolerance properties more suitable for the lower spar cap, could beperformed after the application of the first aging step of either a 2-or 3-step practice in the instantly disclosed alloy. The other alloycould be another 7XXX alloy or quite different in composition, forexample an aluminum-lithium alloy in accordance with U.S. Pat. No.4,961,792, and will have its own typical aging practice which may becomprised of one, two or three steps. Since the post-weld aging of thetwo joined alloy products must necessarily occur together, the agingpractice for the instantly disclosed alloys may need to be two or threesteps depending on the aging requirements of the alloy to which it isjoined. Thus, the flexibility of the instantly disclosed alloys withrespect to the number of aging steps and times that can be successfullyutilized is beneficial for welded multi-alloy components. Even so, somecompromises to the typical aging practice for each alloy may be requireddepending on the specific alloys involved.

The manufacture and aging of a multi-alloy component utilizing theinstantly disclosed alloys joined by welding could be somewhatsimplified by using 7XXX alloys with similar compositions to theinstantly disclosed alloys, but that are leaner or richer in alloyingelements added for strengthening to achieve the desired balance ofstrength and toughness in each component. The typical pre- and post-weldaging practices for such alloys would likely be more compatible than formore dissimilar alloys requiring fewer adjustments to their typicalpractices. Alternatively, the desired differences in strength andtoughness could likely be achieved in some cases with the use of theinstantly disclosed alloys alone by employing different tempers. Forexample, a multi-temper spar solely made from the instantly disclosedalloys could use a the high strength −T79 temper in the upper cap, themoderate strength, higher toughness −T76 temper in the spar web, and thelower strength, highest toughness −T73 temper in the lower spar cap.Typically, the aging times for the −T76 and −T73 would be greater thanfor the −T79 temper. In a welded multi-temper spar, the pre-weld agingfor the −T79 upper spar could be, for example, comprised of a first steponly, the −T76 spar web comprised of a first step and a portion of thesecond step and the −T73 lower spar cap of a first step and a largerportion of the second step. This could be carried out separately on eachcomponent or by staggering their removal from the same furnace. Oncewelded, the same post weld aging practice would be used on the joinedcomponents. With the appropriate selection of the pre- and post-weldaging practice the typical aging practice can be applied to eachcomponent essentially without compromise.

Example 1

Ingots A-D having compositions similar to the embodiments describedabove for the instant alloy family were cast as large commercial scaleingots. In addition, one ingot of aluminum alloy 7085 was cast as acontrol. The ingots were scalped and homogenized with a final soaktemperature of about 870° to about 900° F. One ingot each of alloys Aand B was hot rolled to plate having a thickness of 1.07 inches and awidth of 135 inches. Another ingot each of alloys A and B was hot rolledto a plate having a thickness of 1.10 inches and width of 111 inches.The former will be hereafter referred to as Plate 1 and the latter asPlate 2. One ingot each of alloys C and D was hot rolled to the samethickness and width as Plate 2. Plate 1 and Plate 2 sizes arerepresentative of upper wing panels of an ultra-large capacity aircraft.The 7085 control alloy was hot rolled to the same thickness and width asPlate 1. The plates were solution heat treated between about 880° toabout 895° F. for about 70-100 minutes, water spray quenched to ambienttemperature, and cold stretched about 1.5 to about 3%. Samples from theplates of alloys A thru D and the 7085 control were aged to a highstrength T79-type temper suitable for upper wing components using aconventional three-step aging practice (e.g., as provided by U.S. Pat.No. 6,972,110). The three-step practice consisted of a first step ofabout 6 hours at about 250° F., a second step of about 7 hours at about308° F. and a third step of about 24 hours at about 250° F. In addition,samples of an improved version of aluminum alloy 7055 (U.S. Pat. No.7,097,719) were cut from a number of different production lots of plateof the same or similar width and thickness and given the high strengthT7951 temper and several overaging tempering treatments to decrease thestrength level and increase fracture toughness. The composition of theingots A-D, and the compositions of various conventional alloys areillustrated in Table 2. The aging practice for the −T7951 temper of theimproved version of 7055 was a two-step practice consisting of a firststep of 10 hours at 302° F. and a second step of 6 hours. The overagedtempers were obtained by increasing the first step from about 10 hoursto about 19 to about 24 hours.

TABLE 2 wt. % wt. % wt. % wt. % wt. % wt. % Alloy Zn Cu Mg Fe Si Zr A7.7 1.81 1.62 0.024 0.014 0.11 B 764 2.15 1.65 0.028 0.021 0.10 C 8.052.08 1.78 0.044 0.026 0.12 D 7.83 2.17 1.84 0.036 0.020 0.11 7085 sample7.6 1.62 1.48 0.032 0.015 0.11 7085 AA range 7.0-8.0 1.3-2.0 1.2-1.80.08  0.06  0.08-0.15 max max 7055 Improved 7.6-8.4 2.0-2.6 1.8-2.30.09  0.06  0.08-0.25 max max 7055 AA 7.6-8.4 2.0-2.6 1.8-2.3 0.15 0.10  0.08-0.25 Range max max 7449 AA 7.5-8.7 1.4-2.1 1.8-2.7 0.15 0.12  (1) Range max max (1) 0.25 max Zr + Ti

The tensile and compressive strength, plane strain (K_(Ic)) and apparentplane stress (K_(app)) fracture toughness and exfoliation resistance ofexample alloys A thru D and the 7085 and improved 7055 controls weremeasured. Tensile testing was performed in accordance with testingstandards ASTM E8 and ASTM B557 and compression testing in accordancewith ASTM E9. Plane strain (K_(Ic)) fracture toughness testing wasconducted in accordance with ASTM E399. The plane strain fracturetoughness specimens were of full plate thickness and had a width W of 3inches. Plane stress (K_(app)) fracture toughness testing was conductedin accordance with ASTM E561 and B646. Those skilled in the art willappreciate that the numerical value of K_(app) typically increases asthe test specimen width increases. K_(app) is also influenced byspecimen thickness, initial crack length and test coupon geometry. Thus,K_(app) values can only be reliably compared from test specimens ofequivalent geometry, width, thickness and initial crack length.Accordingly, testing on the example alloys and the 7085 and 7055controls were all performed using center-cracked M(T) specimens havingthe same nominal dimensions, a width of 16 inches, a thickness of 0.25inch and an initial fatigue pre-crack length (2 ao) of 4 inches. Thespecimens were centered at mid-thickness (T/2) of the plate. Exfoliationtesting using the EXCO method was also performed in accordance with ASTMG34. Test specimens were taken at mid-thickness (T/2) and one-tenththickness (T/10).

The measured properties of example alloys A thru D and the nominal 7085composition are given in Table 3. Alloy A exhibited an approximately 3ksi increase in tensile yield and ultimate tensile strength over thenominal 7085 composition in Plate 1 size in both the L and LT direction,a strength increase of about 4%; while alloy B exhibited about a 5 ksiincrease in tensile yield and ultimate tensile, an improvement of about6%. Alloys C and D exhibited even higher strength. The increase in yieldand ultimate tensile strength for both alloys was about 7 ksi, animprovement of about 8%. These are considered significant strengthimprovements by aircraft manufacturers. The improvement in strength wasobtained while retaining excellent exfoliation resistance, all specimensof the example alloys achieving an EA rating.

TABLE 3 UTS TYS CYS KIc Kapp Alloy/Panel Dir (ksi) (ksi) (ksi) (ksi√in)(ksi√in) EXCO 7085 Sample L 83.7 79.9 81.4 50.6 128.9 EA (t/2) LT 83.779.6 na 41.1 102.6 EA (t/10) Example L 86.7 83.2 84.3 50.9 127.5 EA(t/2) Alloy A LT 86.8 82.6 na 40.8 94.0 EA (t/10) Plate 1 Example L 85.881.7 83.0 49.1 129.2 EA (t/2) Alloy A LT 85.7 81.5 na 39.6 91.9 EA(t/10) Plate 2 Example L 89.3 85.7 86.7 43.8 113.2 EA (t/2) Alloy B LT89.2 85.0 na 34.2 78.6 EA (t/10) Plate 1 Example L 87.8 84.3 86.4 43.6129.1 EA (t/2) Alloy B LT 88.5 84.1 na 34.5 86.0 EA (t/10) Plate 2Example L 90.2 87.2 86.5 36.0 115.6 EA (t/2) Alloy C LT 90.2 84.6 na30.0 71.2 EA (t/10) Example L 90.4 87.1 86.2 40.1 107.9 EA (t/2) Alloy DLT 90.6 86.5 na 31.5 68.8 EA (t/10)

The combinations of strength and toughness of example alloys A thru Dare shown in FIGS. 3A, 3B and 4 where they are compared to prior artalloys. FIGS. 3A and 3B compare plane-strain fracture toughness K_(Ic)in the L-T orientation, which corresponds to the principal direction ofloading in the upper wing, as a function of the minimum tensile yieldstrength in the L (rolling) direction of example alloys A thru D, the7085 sample control lots (Table 3), another four lots of 7085 thin plategiven a lower strength aging practice more suitable for lower wings(Table 1), and values from the improved version of 7055 in the T7951temper and with overaged tempering treatments. In addition, typicalfracture K_(Ic) fracture toughness of other prior art alloys in thinplate form are shown. For the example alloys and the overaged tempers of7055, for which no material specifications currently exist, the minimumtensile yield strength was estimated by subtracting 3 ksi from themeasured value. One minimum performance line for the instantly disclosedalloys is designated by the line A-A, which has an equation ofFT=−2.3*(TYS)+229, wherein TYS is the L tensile yield strength of theplate in ksi as measured in accordance with ASTM Standard E8 and ASTMB557, and wherein FT is the L-T plane strain fracture toughness of theplate in ksi√inch as measured in accordance with ASTM E399.

FIG. 3A also includes a shaded region highlighting potential propertiesof thin plate alloy products of the instant disclosure. The shadedregion is bound by a minimum L-T toughness of 36 ksi√inch, a minimumstrength of 74 ksi, and line A-A, which has an equation ofFT=−2.3*(TYS)+229, as provided above. The shaded region of FIG. 3A isparticularly suited for thin plate alloy products of a T74 temper,although alloys having other tempers (e.g., T6, T73, T76, T79) may beproduced that may have properties that lie within the shaded region.

FIG. 3B also includes a shaded region highlighting potential propertiesof thin plate alloy products of the instant disclosure. The shadedregion is bound by a minimum toughness of 30 ksi√inch, a minimumstrength of 79 ksi, and line A-A, which has an equation ofFT=−2.3*(TYS)+229, as provided above. The shaded region of FIG. 3B isparticularly suited for thin plate alloy products of a T76 temper,although alloys having other tempers (e.g., T6, T73, T74, T79) may beproduced that may have properties that lie within the shaded region.

FIG. 4 compares the L tensile yield strength and the apparent planestress fracture toughness (K_(app)) of embodiments of the instantlydisclosed alloys in the L-T orientation again with the five lots of 7085and values from improved 7055. The improved combination of strength andtoughness of 7085 with respect to the improved version of 7055 isobvious. One minimum performance line for the instantly disclosed alloysis designated by the line B-B, which has an equation ofFT=−4.0*(TYS)+453, where TYS is the L tensile yield strength of theplate in ksi as measured in accordance with ASTM Standard E8 and ASTMB557, where FT is the L-T plane stress fracture toughness (K_(app)) ofthe plate in ksi√inch, where FT is measured in accordance with ASTMStandard E561 and B646 on a center-cracked aluminum alloy specimen takenfrom the T/2 location of an aluminum alloy plate, and where the specimenhas a width of 16 inches, a thickness of 0.25 inch and a initial fatiguepre-crack length of 4 inches

Even with significant overaging to achieve the same or similar strengthlevel as in the instantly disclosed alloy, the fracture toughness of7055 is significantly lower. Since the Cu and Mg levels in the instantlydisclosed alloys lies between that of 7085 and the improved version of7055, while Fe and Si levels are similarly low, the expectation was thatthe combination of strength and toughness achievable in the instantlydisclosed alloys would fall between that of 7085 and improved 7055.Surprisingly, the instantly disclosed alloys exhibited an improvedcombination of strength and toughness, not only over improved 7055 butalso over 7085. Thus embodiments of the instantly disclosed alloysidentify a “sweet” composition region which offers higher combinationsof strength and fracture toughness than exhibited by prior art alloys.While the K_(app) values and relative improvements correspond to a testcoupon of the type and dimensions noted, it is expected that similarrelative improvements will be observed in other types and sizes of testcoupons. However, those skilled in the art will also appreciate that theactual K_(app) values may vary significantly in other specimen types andsizes as previously described and the magnitude of the difference mayalso vary.

FIG. 4 also includes a shaded region highlighting potential propertiesof thin plate alloy products of the instant disclosure. The shadedregion is bound by a minimum toughness (K_(app)) of 100 ksi√inch, aminimum tensile yield strength of 80 ksi, and line B-B, which has anequation of FT=−4.0*(TYS)+453, as provided above. The shaded region ofFIG. 4 is particularly suited for thin plate alloy products of a T79temper, although alloys having other tempers (e.g., T6, T73, T74, T76)may be produced that may have properties that lies within the shadedregion. Furthermore, some thin plate products of the instant disclosuremay be able to realize both the plane stress fracture toughness andtensile yield strength values defined by the shaded region of FIG. 4 aswell as the plane strain fracture toughness and tensile yield strengthvalues defined by the shaded region of FIGS. 3A and/or 3B.

Example 2

Four sets of samples in the solution heat treated, quenched andstretched condition (W51 temper) from example alloys A and B platefabricated in Example 1 were given the first two aging steps of thethree-step practice used in Example 1. Subsequently, the first set ofsamples were given a third step with an aging time of 24 hours, the sameas that employed in Example 1, while the second and third set were givenshorter aging times, of 6 and 12 hours. In the fourth set of samples,the third step was not applied (0 hours). Tensile specimens with adiameter of 0.125 inch were machined in the long transverse (LT)direction and short transverse (ST) direction for both an alternateimmersion (AI) stress corrosion cracking resistance test and seacoast(SC) exposure test (also sometimes referred to herein as the seacoastenvironment stress corrosion cracking resistance test). Alternateimmersion testing was conducted in accordance with ASTM G44, G47 and/orG49. More specifically, the specimens were exposed to cycles ofimmersing in a 3.5% NaCl aqueous solution for 10 minutes, followed by 50minutes of air drying while being stressed under a constant strainnecessary to achieve the desired stress level. The seacoast exposuretesting was conducted at Alcoa's Pt. Judith, R.I. seacoast exposuresite, as described below.

Three 3rd step aging times, 0, 12 and 24 hours, and two stress levels,16 and 20 ksi were selected for the ST direction. The first stress levelrepresents the minimum requirement for current upper wing alloys, 7055and 7449 in the ST direction. The second stress level corresponds to a25% higher stress level. The exposure period for AI testing for 7XXXalloys for the ST direction is typically 20 or 30 days or until failureoccurs. In these tests, the maximum exposure period for AI was extendedto 150 days to better assess the performance of the different agingpractices. For seacoast exposures, the maximum exposure period was 466days. The results of the stress corrosion cracking (SCC) tests are givenin Table 4.

TABLE 4 LT Tensile 3rd Step (ksi) SCC Testing Alloy Panel Time (h) YSUTS Location Stress (ksi) # of tests Failure (days) A 2 0 82.5 86.2 AI16 5 48, 101, 101, 101, 115 AI 20 5 32, 59, 70, 101, 115 SC 16 5 297,311 SC 20 5 290, 290, 339, 349 A 2 12 83.8 87.4 AI 16 5 78, 97, 101 AI20 5 53, 98, 101, 101, 101 SC 16 5 325, 339 SC 20 5 66, 325, 339, 367 A2 24 83.7 87.3 AI 16 5 101, 101, 101, 115, 129 AI 20 5 44, 73, 98, 101,143 SC 16 5 332 SC 20 5 332, 346, 346, 402 A 1 12 84.1 87.6 AI 16 5 87,129, 143, 143 AI 20 5 59, 98, 101, 101, 101 SC 16 5 325, 332, 332, 339SC 20 5 325, 332, 339 B 2 12 85.5 89.2 AI 16 5 115, 135, 135 AI 20 5 29,54, 101, 101, 115 SC 16 5 234, 332 SC 20 5 122, 311, 325

The results from example alloy A, Panel 2 with 3rd step aging times of 0(i.e., no third step) 12 and 24 hours indicate there is no significantdifference in the SCC resistance of the instantly disclosed alloys withor without a 3rd aging step or for a shorter or longer 3rd step agingtime. In all cases, the number of days to failure exceeded the standardexposure times of 20 or 30 days for 7XXX alloys for AI SCC at both the16 ksi stress level, the minimum requirement for current upper wingalloys, and at the 25% higher stress level of 20 ksi. The number of daysto failure was also similar for the 3 different aging times. The SCCresistance of the three 3rd step aging times was also similar for theseacoast exposures. Alloy A, Panel 1 and example alloy B, Panel 2 wereevaluated only for the 12-hour 3rd step aging time. Panel 1 is thinnerand wider than Panel 2 and therefore is expected to have a differentgrain aspect ratio and possibly different SCC resistance. The resultsfor alloy A, Panel 1, appeared to be slightly better than those forPanel 2. The results for alloy B, Panel 2 were similar and possiblybetter than for alloy A, Panel 2.

SCC tests in the LT direction were also conducted. For the LT direction,the exposures were interrupted after 30, 47 and 90 days and the exposedspecimens subjected to breaking load testing in accordance with ASTM G139. The percentage retained or residual strength of the exposedspecimen compared to the unexposed tensile strength was determined. Thestress levels for the LT direction were 42 and 63 ksi, corresponding toapproximately 50% and 75% of the LT yield strength of the instantlydisclosed alloys. This test is a means to obtain more quantitativeinformation in a shorter time, and thus is useful for the more SCCresistant LT direction where specimen failures are expected to occur atlonger times, and possibly with greater scatter, than the less SCCresistant ST direction. In one experiment, breaking load tests wereconducted on example alloys A and B given a 3rd step aging practice of0, 6 and 12 hours after an exposure period of 47 days. In a secondexperiment, breaking load tests were conducted on example alloy A and a7055-T7951 control after exposure periods of 30 and 47 days in AI and 90days seacoast exposure at a stress levels corresponding to 50 and 75% ofthe LT yield strength for each alloy. In both experiments, unstressedsamples were also included. The inclusion of unstressed and stressedsamples allows the strength loss resulting from general corrosion andpitting and the loss from SCC to be separated.

The results of the first experiment are shown in FIG. 5, each pointrepresenting an average of 5 specimens. Here, the percentage retainedstrength is the ratio of the strength of an exposed specimen to that ofan unexposed specimen (i.e., uncorroded) expressed on a percentagebasis. The results indicate there was no loss in general corrosionresistance (unstressed) or SCC resistance (stressed) with theelimination of the 3rd aging step. In fact, the specimens without the3rd step had a greater retained or residual strength than those with a 6or 12-hour 3rd step. For a given aging time, alloy B outperformed alloyA. The results of the second experiment are given in FIG. 6, each pointrepresenting an average of 5 specimens. FIG. 6 is a graph comparing thepercentage retained strength in the LT direction of the instantlydisclosed alloys and prior art alloy 7055 for a 12-hour 2nd step agingtime following exposures of 30 and 47 days in 3.5% NaCl solution and 90days seacoast exposure at stress levels of 50 and 75% of the yieldstrength of each alloy. Example alloy A had greater percentage retainedstrength than the 7055 alloy for all three exposures in both theunstressed and stressed condition and at the two stress levels.

Overall, the corrosion results indicate that both the 2- and 3-stepaging practice provide acceptable corrosion performance of the instantlydisclosed alloys for upper wing applications. One disadvantage of the2-step practice is that the strength is slightly lower as illustrated inTable 4 for example alloy A. Compared to a 3rd step age time of 24hours, the yield strength without the 3rd step was about 1 ksi higher.As previously described, the flexibility in aging practice of theinstantly disclosed alloys is a beneficial characteristic. A 2-steppractice is typical for applications such as upper wing skin andstringer, where the aging is partly or fully applied during an ageforming process by the aircraft manufacturer or subcontractor and it isdesirable that the age forming cycle be as short as practicable tomaximize production throughput. In this regard, the instantly disclosedalloys with the 2-step practice utilized herein, which had a total soaktime of 13 hours, offers an improvement over the current upper wingalloys. Depending on the age forming requirements, this could possiblybe shortened further to about 7 hours if the first step is applied bythe material producer and only the second step is carried out in the ageforming process.

A 3-step practice may be used when the material is supplied by theproducer in the fully aged condition for applications such as an upperwing spar or spar web in a built-up design. A lower strength temper,such as a T76 or T74 temper, using either a 2- or 3-step practice, mayalso be used for these applications depending on the requirements andthe direction of the design stresses relative to the alloy productsgrain orientation. When the instantly disclosed alloys are to be weldedto another alloy product and post weld aged as part of a multi-alloycomponent, a 2- or 3-step practice could be used depending on the agingpractice of the alloy or alloys to which the instantly disclosed alloysare to be joined. The flexibility afforded by the instantly disclosedalloys may also be useful for combining the curing cycles of adhesivesused to attach reinforcing materials with the aging of the instantlydisclosed alloys.

Example 3

Samples of example alloy A plate in the solution heat treated, quenchedand stretched condition (W51 temper) fabricated in Example 1 weremachined into panels 0.5 inch thick by 6 inches wide by 35 inches long.Samples from 2099 extrusion were acquired in the T3511 temper andmachined to the same dimensions. In both cases the length dimension wasin the rolling direction 2099 is a commercially availablealuminum-lithium alloy registered with the Aluminum Association havingthe composition 2.4-3.0 wt. % Cu, 0.1-0.5 wt. % Mg, 0.4-1.0 Zn, 0.1-0.5Mn, 0.05-0.12 Zr and 1.6-2.0 Li, the remainder Al and incidentalimpurities. Panels of the example alloy A and 2099 were joined byfriction stir welding with the weld line along the length of the panels.This combination of the instantly disclosed alloys and 2099, which havevery dissimilar compositions could be used, for example, for amulti-alloy spar or rib. In a spar, the instantly disclosed alloys couldbe used in the upper cap and web where high compressive strength isneeded and 2099 in the lower spar cap where high resistance to fatiguecrack growth is beneficial. Similarly, in the rib, the instantlydisclosed alloys could be used in the feet where high strength isimportant and 2099 in the spar web where high stiffness and low densityare beneficial.

Prior to the friction stir welding operation, the alloy A and 2099panels were aged separately. The pre-weld aging for alloy A consisted ofa first step of 6 hours at 250° F. while the pre-weld aging practice of2099 consisted of a first and second step of different times and/ortemperatures than that used for the instantly disclosed alloys. Thepost-weld aging practice of the joined panels was necessarily the sameand consisted of a first step of 6 hours at 250° F. and a second step of18 hours at 305° F. Post-weld aging is desirable for improving thestrength and corrosion properties of the weld area. In order to increasethe weld properties, in particular the strength and corrosionresistance, as much of the aging as possible should be conducted afterwelding. However, for dissimilar alloys the ability to do so may belimited by the individual aging requirements for each alloy and thefinal desired temper for each. The pre-weld aging practices for eachalloy and the post-weld aging practice for the multi-alloy panel werecarefully selected to target a −T76 type temper in the instantlydisclosed alloys and a −T83 type temper in the 2099. Even so, somecompromise in the aging practices of both alloys was necessary and theflexibility of the instantly disclosed alloys with respect to the numberof aging steps and times that can be successfully utilized whileobtaining good properties was beneficial in that regard.

Mechanical properties including tensile strength, compressive strength,tensile and compressive elastic modulus and fracture toughness weremeasured in the base metals (i.e., outside the weld and heat affectedzone), the heat affected zone (HAZ) and the weld following the post-weldage. The extent of the each region and the position of the specimenstherein were determined using Vicker's micro-hardness (VHN) measurementsacross the weld and optical micrographs. The testing was performed inaccordance with the applicable ASTM test methods: ASTM E8 and B557 fortensile testing, E9 for compression testing, E111 for tensile andcompressive modulus testing, and ASTM E399 for plane strain fracturetoughness. Tensile properties were measured in the L and LT directions.Compressive strength and elastic modulus were measured in the Ldirection only. The plane strain fracture toughness specimens were inthe T-L orientation, had a width W of 2 inches and were of full panelthickness. The fracture specimens were excised from the panel so thattheir machined slot (representing the expected plane of crack extension)was aligned with the region of interest. Two specimens were taken in theweld and HAZ, one specimen with the machined notch pointed in the samedirection as the friction stir welding tool had traveled during thewelding operation and one with the machined notch pointed in theopposite direction. The results from these tests are given in Table 5.

TABLE 5 Alloy A 2099 Property Dir Base Metal HAZ Weld HAZ Base Metal UTS(ksi) L 84.5 56.8 61.2 77.6 83.1 LT 84.3  62.4* 77.2 TYS (ksi) L 79.843.9 59.0 69.9 76.0 LT 79.1  50.3* 70.6 CYS (ksi) L 82.3 69.5 60.9 76.2Et (Mpsi) L 10.3 10.4 11.5 11.5 Ec (Mpsi) L 10.7 10.7 11.3 11.8 11.9KIc, Kq (ksi√in) T-L 41.5 34.2¹, 36.2² 40.5¹, 38.2² 26.4¹, 27.1² 32.1Notes: *LT tensile specimens traversed both the weld and HAZ failing inthe weakest location. ¹Crack extension in the same direction as thattraveled by the welding tool during the welding operation. ²Crackextension in the opposite direction to that traveled by the welding toolduring the welding operation.

Even with the compromise in aging practices made for each alloy, thebase metal of each which received the pre-weld aging practice (differentfor each alloy) and post-weld aging practice (the same for each alloy)achieved the desired strength and toughness level for the targetedtempers. The properties in the HAZ and weld were lower as is typicallyobserved for welds. The weld region is essentially solution heat treatedduring the friction stir welding process so the artificial aging of thisregion occurs only during the post-weld age. Likewise, the HAZ is alsoheated during the welding process but at a temperature which is belowthat used for solution heat treating and thus inadequate to fullysolutionize the alloying elements. This can limit its aging response inthe HAZ during the post-weld age and degrade its strength and fracturetoughness. Despite these factors, the weld efficiency (i.e, the ratio ofthe weld strength to the base metal strength) achieved was quite good.Measured perpendicular to the weld line where the tensile specimenincluded both the weld and HAZ, the weld efficiency was 71% for tensileyield strength (TYS) and 81% for ultimate tensile strength compared tothe base metal strength for 2099 in the LT direction.

The fracture toughness achieved in the weld and HAZ were alsosatisfactory. In the weld zone the fracture toughness was equivalent tothat in the alloy A base metal, while the fracture toughness in the HAZon both the alloy A and 2099 side of the weld were lower than in theirrespective base metal but still sufficient to meet the requirements ofmost aircraft structure.

Stress corrosion cracking (SCC) and exfoliation testing were alsoperformed on the joined panels following the post-weld age. For SCCtesting, flat tensile-type specimens with a thickness of 0.235 inch weremachined at mid-thickness perpendicular to and across the weld and HAZ.Three specimens each were tested at two stress levels, 26 and 35 ksi byalternate immersion in accordance with ASTM G44, G47 and/or G49. Nofailures were observed after an exposure period of 250 days. Forexfoliation testing, two rectangular specimens of full panel thicknesscontaining the weld, HAZ and base metals were tested using the EXCO testmethod in accordance with ASTM G34. This test method is an appropriateaccelerated test method for 7XXX alloys such as the instantly disclosedalloys. A second set of specimens of full panel thickness were testedusing Dry Bottom MASTMAASIS in accordance with ASTM G85. This testmethod is an appropriate accelerated test method for 2099. Both the basemetal of alloy A and 2099 had an exfoliation rating of EA. This ratingis indicative of good corrosion performance and consistent with thetypical performance of the targeted tempers for each alloy. The weldregion which contained a mixture of both alloys had a rating of EB bythe EXCO test method, again indicating reasonably good exfoliationcorrosion resistance. Some degradation in corrosion performance of theweld is expected since this region receives only the post-weld age. TheHAZ in 2099 had a MASTMAASIS rating of P, however the HAZ in alloy Aexhibited localized attack and had an EXCO rating of ED. This corrosionperformance may be unacceptable for internal aircraft structure such asspars and ribs but could likely be improved by optimizing the frictionstir welding parameters or using cooling methods during welding in orderto reduce heat input into the HAZ. This region could also be protectedin service by the use of corrosion protection methods. For example,prior to the application of anodize and an anti-corrosive primer, whichare already commonly used for corrosion protection in internalstructure, an aluminum alloy more anodic than the instantly disclosedalloys could be applied along the weld line by thermal spray or othermethods. Galvanic corrosion resulting from differences in corrosionpotential in alloy A and 2099 may have contributed to the localizedattack in the HAZ of alloy. In this case, the use of leaner and richeralloys of similar composition to the instantly disclosed alloys, whichshould have less difference in corrosion potential than two verydissimilar alloys, or the use of the instantly disclosed alloys alone indifferent tempers may be beneficial for improving corrosion resistancein the HAZ.

Example 4

Two ingots are cast as large commercial scale ingots. The ingots have acomposition consistent with the teachings of the instant disclosure. Thefirst ingot is designated alloy E and the second ingot is designatedalloy F. In addition, four ingots of Aluminum Association alloy 7085 andsix ingots of Aluminum Association alloy 7050 are cast as control. Thecomposition of the alloys E and F, the 7050 and 7085 control ingots, andthe composition ranges for 7085 and 7050 registered with the AluminumAssociation are provided in Table 6.

TABLE 6 Alloy wt. % Zn wt. % Cu wt. % Mg wt. % Fe wt. % Si wt. % Zr E7.57 2.11 1.63 0.04 0.01 0.11 F 7.64 2.15 1.65 0.03 0.02 0.1  7050-lot 16.07 2.21 2.18 0.08 0.05 0.11 7050-lot 2 6.07 2.21 2.18 0.08 0.05 0.117050-lot 3 6.00 2.22 2.15 0.08 0.05 0.11 7050-lot 4 6.04 2.29 2.17 0.070.04 0.11 7050-lot 5 6.04 2.29 2.17 0.07 0.04 0.11 7050-lot 6 6.09 2.262.20 0.08 0.04 0.11 7085-lot 1 7.47 1.64 1.50 0.05 0.02 0.11 7085-lot 27.48 1.68 1.50 0.05 0.01 0.11 7085-lot 3 7.35 1.65 1.50 0.04 0.02 0.127085-lot 4 7.31 1.65 1.44 0.03 0.02 0.12 AA7085 range 7.0-8.0 1.3-2.01.2-1.8 0.08 max 0.06 max 0.08-0.15 AA7050 range 5.7-6.7 2.0-2.6 1.9-2.60.15 max 0.12 max 0.08-0.15

The ingots are scalped and homogenized with a final soak temperature ofabout 870° to 910° F. The ingot with composition E is hot rolled toplate having a thickness of 3.125 inches while ingot with composition Fis hot rolled to plate having a thickness of 4.0 inches. Such dimensionsare representative of standard aerospace plate used for integrallymachined parts. Lots 1-3 of the 7085 control ingots are hot rolled toplate having a thickness of about 4 inches. Lot 4 of the 7085 controlingot is hot rolled to plate having a thickness of about 3 inches. Three7050 control ingots are hot rolled to plate having a thickness of about4 inches. Another three 7050 control ingots are hot rolled to platehaving a thickness of about 3 inches. All ingots were cross-rolled inthe long transverse direction by less than 15%. All plates were solutionheat treated between about 880° and 900° F. for about 2 to 4 h, waterspray quenched to ambient temperature, and cold stretched about 1.5 to3%.

Samples from the alloy E and F plates were obtained. These samples wereaged to a T74-type temper (suitable for integrally machined components)using a conventional three-step practice. The three-step practiceconsisted of a first step of about 6 hours at 250° F., a second stepbetween 15 and 20 h at a temperature of 310° F. and a third step ofabout 24 hours at 250° F. Some of the alloy E and F samples were agedfor 15 hours during the second step (samples 1). Others of the alloy Fsamples were aged for 18 hours during the second step (sample 2). Othersof the alloy E samples were aged for 20 hours during the second step(sample 2). The 7085 fl-inch control lots were also aged to a T74 temperusing this conventional 3-step aging process. Sample 1 of lot 4 (3-inchplate) of the 7085 control lots was aged to a T76 temper using aconventional 3-step aging process, and sample 2 of lot 4 (3-inch plate)of the 7085 control lots was aged to a T74 temper using a conventional3-step aging process. The 7050 control lots were aged to a T74 temperusing a conventional 2-step aging process.

The tensile properties and plane strain (K_(Ic)) fracture toughness ofthe samples of alloys E and F and the 7085 and 7050 control lots weremeasured. Tensile testing was performed in accordance with ASTM E8 andASTM B557. Plane strain (K_(Ic)) fracture toughness testing wasperformed in accordance with ASTM E399. The plane strain fracturetoughness specimens for alloy E were 2 inches thick and had a width W of4 inches in the T-L orientation, and were 1 inch thick and had a width Wof 2 inches in the S-L orientation. The plane strain fracture toughnessspecimens for alloy F were 1 inch thick and had a width W of 2 inches inboth the T-L and S-L orientations. The fracture toughness specimens foralloys E and F were centered at mid-thickness (T/2) of the plate. Theplane strain fracture toughness specimens for the 4-inch control 7085plates were 2 inches thick and had a width W of 4 inches in the T-Lorientation, and were 1.5 inches thick and had a width W of 3 inches inthe S-L orientation. The plane strain fracture toughness specimens forthe 3-inch control 7085 plates were 1.75 inch thick and had a width W of5 inches in the T-L orientation, and were 1.25 inches thick and had awidth W of 2.5 inches in the S-L orientation. The fracture toughnessspecimens for the 4-inch control 7085 plates were centered atquarter-thickness (T/4) of the plate in the T-L orientation and atmid-thickness (T/2) of the plate in the S-L orientation. The fracturetoughness specimens for the 3-inch control 7085 plates were centered atmid-thickness (T/2) of the plate in both the T-L and S-L orientations.The plane strain fracture toughness specimens in the T-L orientation forthe control 7050 plates were 2 inches thick and had a width W of 4inches. The plane strain fracture toughness specimens in the S-Lorientation for the 3-inch thick control 7050 plates were 1 inch thickand had a width W of 2 inches. The plane strain fracture toughnessspecimens in the S-L orientation for the 4-inch thick control 7050plates were 1.5 inches thick and had a width W of 3 inches. The fracturetoughness specimens for the control 7050 plates were centered atmid-thickness (T/2) of the plate in both the T-L and S-L orientations.Exfoliation testing using the EXCO method was performed for alloy F inaccordance with the ASTM G34 standard, where test specimens were takenat mid-thickness (T/2), quarter-thickness (T/4) and one-tenth thickness(T/10).

The measured properties of alloys E and F and the 7085 and 7050 controllots are provided in Table 7. At a plate thickness of about 3 inches,alloy E exhibited an approximately 9 to 12 ksi increase in tensile yieldstrength and about a 6 to 8 ksi increase in ultimate tensile strengthover the 7050 control lots in the LT direction. Similarly, alloy Eexhibited an approximately 8 to 10 ksi increase in tensile yieldstrength and about a 6 to 8 ksi increase in ultimate tensile strengthover the 7050 control lots in the ST direction. At a plate thickness of4 inches, alloy F exhibited an approximately 7 to 9 ksi increase intensile yield strength and about a 3 to 4 ksi increase in ultimatetensile strength over the 7050 control lots in the LT direction.Similarly, alloy F exhibited an approximately 5 to 7 ksi increase intensile yield strength and about a 4 to 5 ksi increase in ultimatetensile strength over the 7050 control lots in the ST direction. Alloy Fexhibits strength improvements of about 2 to 5 ksi for tensile yield andultimate strength in both the LT and ST directions compared to the 7085control lots of a T74 temper. These strength improvements are consideredsignificant strength improvements by aircraft manufacturers.

TABLE 7 Lot/ Sample Thickness TYS UTS Elongation KIc Alloy No (inch)Direction (ksi) (ksi) (%) Orientation (ksiVin) 7050- Lot 1 3 LT 66.276.6 11.4 T-L 28.2 T7451 ST 62.0 73.4 6.2 S-L 28.0 Lot 2 3 LT 65.8 76.211.4 T-L 29.2 ST 61.7 73.3 6.7 S-L 28.1 Lot 3 3 LT 65.3 75.3 11.0 T-L30.0 ST 61.0 72.5 6.6 S-L 28.8 Lot 4 4 LT 65.2 75.8 11.3 T-L 26.3 ST62.9 74.6 5.8 S-L 22.4 Lot 5 4 LT 65.6 76.1 10.8 T-L 26.4 ST 62.4 73.55.6 S-L 26.6 Lot 6 4 LT 66.9 76.8 7.9 T-L 26.5 ST 61.6 73.0 5.1 S-L 26.27085- Lot 1 4 LT 69.1 76.4 10.5 T-L 29.1 T7451 ST 64.4 74.1 7.5 S-L 32.3Lot 2 4 LT 69.9 76.5 10.7 T-L 29.4 ST 64.7 74.3 7.0 S-L 31.0 Lot 3 4 LT69.5 76.9 10.6 T-L 30.1 ST 65.4 74.8 6.2 S-L 32.1 7085- Lot 4, 3 LT 69.375.4 18.2 T-L 35.4 T7X51 sample 2 3 ST 66.5 75.0 13.5 S-L 39.6 Lot 4, 3LT 68.6 74.5 19.0 T-L 37.2 sample 1 3 ST 65.5 74.1 13.9 S-L 37.7 Alloy ESample 1 3.125 LT 77.6 83.8 9.3 T-L 25.0 ST 71.5 80.9 7.8 S-L 27.6Sample 2 3.125 LT 74.7 82.0 9.7 T-L 26.9 ST 69.7 79.4 8.6 S-L 29.4 AlloyF Sample 1 4 LT 74.5 80.3 10.0 T-L 26.4 ST 69.2 78.2 7.8 S-L 25.1 Sample2 4 LT 73.0 79.6 10.0 T-L 28.3 ST 67.3 77.6 8.6 S-L 27.4

The properties of alloy E and various conventional alloys having athickness of about 3 inches are illustrated in FIG. 7. Moreparticularly, FIG. 7 compares the plane-strain fracture toughness(K_(Ic)) in the T-L orientation as a function of the tensile yieldstrength in the LT (long transverse) direction for alloy E (thickness of3.125 inches), the 7050 control lots (having a thickness of about 3inches) and data from the 3-inch 7085 lot. Alloy E realizes asignificantly higher tensile yield strength with similar toughness tothe 7050 control lots. Alloy E also realizes a strength-to-toughnessrelationship comparable to the 7085 alloy, but, as described below, the7085 alloy is unable to consistently pass a seacoast environment SCCtest. In other words, alloy E realizes an equal or better stresscorrosion resistance than a similarly produced and sized 7085 alloy, butat a higher LT strength. Thus, alloy E realizes a heretofore unrealizedcombination of LT strength, T-L toughness, and corrosion resistance atthe stated thickness range.

FIG. 7 also includes a shaded region highlighting potential propertiesof alloy plate products of the instant disclosure. The shaded region isbound by a minimum toughness of 22 ksi√inch, a minimum strength of 72ksi, and line C-C, which has an equation of FT_TL=−1.0*(TYS_LT)+98,where TYS_LT is the LT tensile yield strength of the plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, and whereFT_TL is the T-L plane strain fracture toughness of the plate inksi√inch as measured in accordance with ASTM E399. The shaded region ofFIG. 7 is particularly suited for plate alloy products having athickness of from about 2.0 to 2.5 inches to about 3.0, 3.125 or 3.25inches, and of a T73, T74, T76 or T79 temper.

The properties of alloy F and various conventional alloys having athickness of 4 inches are illustrated in FIG. 8. More particularly, FIG.8 compares the plane-strain fracture toughness (K_(Ic)) in the T-Lorientation as a function of the tensile yield strength in the LT (longtransverse) direction for alloy F (thickness of 4.0 inches), the 7050control lots (having a thickness of about 4 inches) and the 4-inch 7085control lots. Alloy F realizes a significantly higher tensile yieldstrength with similar toughness to the 7050 control lots. Alloy F alsorealizes a strength-to-toughness relationship similar to that of the7085 control lots, but, as described below, the 7085 alloy is unable toconsistently pass a seacoast environment SCC test. In other words, alloyF achieves an equal or better stress corrosion resistance than asimilarly produced and sized 7085 alloy, but at a higher LT strength.Thus, alloy F realizes a heretofore unrealized combination of LTstrength, T-L toughness, and corrosion resistance at the statedthickness range.

FIG. 8 also includes a shaded region highlighting potential propertiesof alloy plate products of the instant disclosure. The shaded region isbound by a minimum toughness of 21 ksi√inch, a minimum strength of 71ksi, and line D-D, which has an equation of FT_TL=−1.0*(TYS_LT)+98,where TYS_LT is the LT tensile yield strength of the plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, where FT_TLis the T-L plane strain fracture toughness of the plate in ksi√inch asmeasured in accordance with ASTM E399. The shaded region of FIG. 8 isparticularly suited for plate alloy products of having a thickness offrom about 3.0 to 3.125 or 3.25 inches to about 3.5, 3.75 or 4 inches,and of a T73, T74, T76 or T79 temper.

The properties of alloy E and various conventional alloys having athickness of about 3 inches are also illustrated in FIG. 9. Moreparticularly, FIG. 9 compares the plane-strain fracture toughness(K_(Ic)) in the S-L orientation as a function of the tensile yieldstrength in the ST (short transverse) direction for alloy E (thicknessof 3.125 inches) and the 7050 control lots (having a thickness of about3 inches) and the 3-inch 7085 control lot. Alloy E realizes asignificantly higher tensile yield strength with similar toughness tothe 7050 control lots. Alloy E also realizes a strength-to-toughnessrelationship similar to that of the 7085 control lot, but, as describedbelow, the 7085 alloy is unable to consistently pass a seacoastenvironment SCC test. In other words, alloy E realizes an equal orbetter stress corrosion resistance than a similarly produced and sized7085 alloy, but at a higher ST strength. Thus, alloy E realizes aheretofore unrealized combination of ST strength, S-L toughness, andcorrosion resistance at the stated thickness range.

FIG. 9 also includes a shaded region highlighting potential propertiesof alloy plate products of the instant disclosure, The shaded region isbound by a minimum toughness of 22 ksi√inch, a minimum strength of 69ksi, and line E-E, which has an equation of FT_SL=−1.1*(TYS_ST)+99,where TYS_ST is the ST tensile yield strength of the plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, whereinFT_SL is the S-L plane strain fracture toughness of the plate inksi√inch as measured in accordance with ASTM E399. The shaded region ofFIG. 9 is particularly suited for plate alloy products having athickness of from about 2.0 to 2.5 inches to about 3.0, 3.125 or 3.25inches, and of a T73, T74, T76 or T79 temper.

The properties of alloy F and various conventional alloys having athickness of about 4 inches are also illustrated in FIG. 10. Moreparticularly, FIG. 10 compares the plane-strain fracture toughness(K_(Ic)) in the S-L orientation as a function of the tensile yieldstrength in the ST (short transverse) direction for alloy F (thicknessof 4.0 inches) and the 7050 control lots (having a thickness of about 4inches), and the 7085 control lots (having a thickness of about 4inches). Alloy F realizes a significantly higher tensile yield strengthwith similar toughness to the 7050 control lots. Alloy F also realizes astrength-to-toughness relationship similar to that of the 7085 controllots, but, as described below, the 7085 alloy is unable to consistentlypass a seacoast environment SCC test. In other words, alloy F achievesan equal or better stress corrosion resistance than a similarly producedand sized 7085 alloy, but at a higher ST strength. Thus, alloy Frealizes a heretofore unrealized combination of ST strength, S-Ltoughness, and corrosion resistance at the stated thickness.

FIG. 10 also includes a shaded region highlighting potential propertiesof alloy plate products of the instant disclosure. The shaded region isbound by a minimum toughness of 20 ksi√inch, a minimum strength of 66ksi, and line F-F, which has an equation of FT_SL=−1.1*(TYS_ST)+99,where TYS_ST is the ST tensile yield strength of the plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, whereinFT_SL is the S-L plane strain fracture toughness of the plate inksi√inch as measured in accordance with ASTM E399. The shaded region ofFIG. 10 is particularly suited for plate alloy products having athickness of from about 2.0 to 2.5 inches to about 3.0, 3.125 or 3.25inches, and of a T73, T74, T76 or T79 temper.

Alloy F, in both aging conditions, was tested for exfoliation corrosionresistance (EXCO) in accordance with ASTM G34. Alloy F in both agingconditions achieved an EA rating consistent with good exfoliationcorrosion resistance for a 7XXX alloy. Similar results would be expectedif alloy E was subjected to EXCO testing. Thus, the instantly disclosedalloys obtain improvements in strength while retaining excellentexfoliation resistance characteristic, all specimens of alloy Fachieving an EXCO rating of EA.

Alloys E, F and 7085 alloys were subjected to two types of stresscorrosion cracking tests. A first test, an alternate immersion (AI)accelerated stress corrosion cracking (SCC) testing was performed forsamples 1 and 2 of alloys E and F, as well as on the 7085 control plateson test specimens taken at mid-thickness (T/2) in the ST direction inaccordance with ASTM G44, G47 and/or G49 standards. AI SCC testingresults are illustrated in Table 8 (4-inch) and Table 9 (3-inch).

TABLE 8 Lot/ Thick- # Days TYS Sample ness Stress of in Failures Alloy(LT) No (inch) (ksi) tests test (days) Alloy F 74.5 ksi Sample 1 4 40 3100 53, 71, 78 50 3 100 51, 57, 58 73.0 ksi Sample 2 4 40 3 100 74, 93,100 50 3 100 63, 63, 63 7085- 69.1 ksi Lot 1 4 35 5 65 No T7451 failures45 5 65 44, 50 and 58 69.9 ksi Lot 2 4 35 5 65 60, 60 and 62 45 5 65 45,46, 57, 57, 57 69.5 ksi Lot 3 4 35 5 65 No failures 45 5 65 50, 57

TABLE 9 Lot/ Thick- # Days TYS Sample ness Stress of in Failures Alloy(LT) No (inch) (ksi) tests test (days) Alloy E 77.6 ksi Sample 1 3.12540 3 100 54, 54, 99 50 3 100 47, 52, 74 74.7 ksi Sample 2 3.125 40 3 10068, 74, 76 50 3 100 50, 54, 57 7085- 68.6 ksi Lot 4, 3 35 5 100 72, 73,75 T7451 sample 2 45 5 100 57, 61, 61, 64, 65 7085- 69.4 ksi Lot 4, 3 355 100 61, 65, 68, T7651 sample 1 81 45 5 100 48, 65, 65

Alloys E and F realized acceptable performance at stress levels of 40and 50 ksi, which is 5 and 15 ksi, respectively, above the minimumrequirements for qualifying an alloy as having a T74 temper.

Seacoast environment SCC testing was also performed for samples 1 and 2of alloy E on test specimens taken at mid-thickness (T/2) in the STdirection. Seacoast environment SCC testing for alloy 7085 was alsoobtained. The specimens for the seacoast environment SCC testing aretested in constant strain fixtures (e.g., similar to those use inaccelerated laboratory SCC testing). The seacoast SCC testing conditionsinclude continuously exposing the samples via racks to a seacoastenvironment, where the samples are about 1.5 meters from the ground, thesamples are oriented 45° from the horizontal, and a face of the sampleface the prevailing winds. The samples are located about 100 meters fromthe coastline. In one embodiment, the coastline is of a rocky nature,with the prevailing winds oriented toward the samples so as to providean aggressive salt-mist exposure (e.g., a location similar to theseacoast exposure station, Pt. Judith, R.I., USA of Alcoa Inc.).Seacoast SCC testing results for alloy E and 7085 alloys are illustratedin Table 10.

TABLE 10 Lot/ Thick- # Days TYS Sample ness Stress of in Failures Alloy(LT) No (inch) (ksi) tests test (days) Alloy E 77.6 ksi Sample 1 3.12540 3 262 No failures 50 3 262 60, 102 74.7 ksi Sample 2 3.125 40 3 262No failures 50 3 262 No failures 7085 68.6 ksi Lot 4, 3 35 5 525 Nosample 2 failures 7085 69.4 ksi Lot 4, 3 35 5 525 76, 132 sample 1 and 3with no failures

Many samples of alloy E had not failed (a specimen fails when itseparates into two pieces or a crack becomes visible with the naked eye)at a stress levels of 40 ksi and 50 ksi after 262 days of exposure.Recall that alloy E achieved LT strengths of 74.7 ksi and 77.6 ksi forsamples 1 and 2, respectively. Conversely, 7085 alloys of similarthickness and only having LT strengths of 68.6, and 69.4 ksi failed zeroout of 5 and 2 out of 5 times respectively. Note the trend in the 7085data that for only minor increases in strength of the 7085 alloy, theability to pass a seacoast environment SCC test decreases. It isanticipated that, if a 7085 alloy were processed to achieve an LTstrength level of 72 ksi at a thickness of 3 inches, such a 7085 alloywould consistently fail a seacoast environment SCC test (at a stress of35 ksi in the ST direction), whereas alloy E (and other alloys definedby the instant disclosure) would consistently pass a seacoastenvironment SCC test at the same strength and SCC stress level.

Thus, the instantly disclosed alloys are able to achieve a heretoforeunrealized combination of strength, toughness, and corrosion resistanceat the stated thickness ranges. In one embodiment, an aluminum alloyproduct of a T74 temper is provided. The aluminum alloy product may bemade from a first plate, a second plate, and/or a third plate. If afirst plate is utilized, the first plate will have a thickness of notgreater than about 2.00 inches, and comprises an alloy composition ofany of embodiments 1, 2, 3, 4 or 5 of FIGS. 2C-1, 2C-2, 2D-1, and 2D-2or embodiments 1 or 2 of FIGS. 2E and 2F, as described above. If asecond plate is utilized, the second plate will have a thickness ofgreater than 2.00 inches, but not greater than 3.00 inches, andcomprises an alloy composition of any of embodiments 1, 2, 3, 4 or 5 ofFIGS. 2C-1, 2C-2, 2D-1, and 2D-2 or embodiments 1 or 2 of FIGS. 2E and2F, as described above. If a third plate is utilized, the third platewill have a thickness of greater than 3.00 inches, but not greater than4.00 inches, and comprises an alloy composition of any of embodiments 1,2, 3, 4 or 5 of FIGS. 2C-1, 2C-2, 2D-1, and 2D-2 or embodiments 1 or 2of FIGS. 2E and 2F, as described above. The aluminum alloy product maycomprise other compositions, such others of the above-noted compositionlevels. Furthermore, in any of these embodiments, the aluminum alloy mayconsist essentially of the stated ingredients (aside from aluminum), thebalance being aluminum and incidental elements and impurities.

In this embodiment, any first plate may have a strength-to-toughnessrelationship that satisfies the expression of FT≧−2.3*(TYS)+229, whereinTYS is the L tensile yield strength of the first plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, wherein FTis the L-T plane strain fracture toughness of the first plate inksi√inch as measured in accordance with ASTM E399, where the first platehas a TYS of at least 74 ksi, where the first plate has a FT of at least36 ksi√inch. In some of these embodiments, the plate may have a tensileyield strength of at least about 75 ksi, such as at least about 76 ksior at least about 77 ksi or at least about 78 ksi or at least about 79ksi, or even at least about 80 ksi. In some of these embodiments, theplate may have a toughness of at least about 40 ksi√inch, such as atleast about 42 ksi√inch or at least about 44 ksi√inch or at least about46 ksi√inch or at least about 48 ksi√inch or even at least about 50ksi√inch.

In this embodiment, any second plate may have a strength-to-toughnessrelationship that satisfies the expression FT_TL≧−1.0*(TYS_LT)+98 whereTYS_LT is the LT tensile yield strength of the second plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, whereinFT_TL is the T-L plane strain fracture toughness of the second plate inksi√inch as measured in accordance with ASTM E399, wherein the secondplate has a TYS_LT of at least 72 ksi, and where the second plate has aFT_TL of at least 24.5 ksi√inch. In some of these embodiments, the platemay have a tensile yield strength of at least about 73 ksi, such as atleast about 74 ksi or as at least about 75 ksi or at least about 76 ksior even at least about 77 ksi. In some of these embodiments, the platemay have a toughness of at least about 25 ksi√inch, such as at leastabout 26 ksi√inch or at least about 27 ksi√inch or even at least about28 ksi√inch.

In this embodiment, any second plate may have a strength-to-toughnessrelationship that satisfies the expression FT_SL≧−1.1*(TYS_ST)+99, whereTYS_ST is the ST tensile yield strength of the second plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, whereinFT_SL is the S-L plane strain fracture toughness of the second plate inksi√inch as measured in accordance with ASTM E399, where the secondplate has a TYS_ST of at least 69 ksi, where the second plate has aFT_SL of at least 25 ksi√inch. In some of these embodiments, the platemay have a tensile yield strength of at least about 69.5 ksi, such as atleast about 70 ksi or as at least about 70.5 ksi or even at least about71 ksi. In some of these embodiments, the plate may have a toughness ofat least about 26 ksi√inch, such as at least about 27 ksi√inch or atleast about 28 ksi√inch or at least about 29 ksi√inch or at least about30 ksi√inch or even at least about 31 ksi√inch.

In this embodiment, any third plate may have a strength-to-toughnessrelationship that satisfies the expression FT_TL≧−1.0*(TYS_LT)+98, whereTYS_LT is the LT tensile yield strength of the third plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, whereinFT_TL is the T-L plane strain fracture toughness respectively of thethird plate in ksi√inch as measured in accordance with ASTM E399, wherethe third plate has a TYS_LT of at least 71 ksi, and where the thirdplate has a FT_TL of at least 23 ksi√inch. In some of these embodiments,the plate may have a tensile yield strength of at least about 71.5 ksi,such as at least about 72 ksi or as at least about 72.5 ksi or at leastabout 73 ksi or at least about 73.5 ksi or even at least about 74 ksi.In some of these embodiments, the plate may have a toughness of at leastabout 24 ksi√inch, such as at least about 25 ksi√inch or at least about26 ksi√inch or at least about 27 ksi√inch or at least about 28 ksi√inchor even at least about 29 ksi√inch.

In this embodiment, any third plate may have a strength-to-toughnessrelationship that satisfies the expression FT_SL≧−1.1*(TYS_ST)+99, whereTYS_ST is the ST tensile yield strength of the third plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, where FT_SLis the S-L plane strain fracture toughness of the third plate inksi√inch as measured in accordance with ASTM E399, where the third platehas a TYS_ST of at least 66 ksi, and where the third plate has a FT_SLof at least 23 ksi√inch. In some of these embodiments, the plate mayhave a tensile yield strength of at least about 66.5 ksi, such as atleast about 67 ksi or as at least about 67.5 ksi or at least about 68ksi or at least about 68.5 ksi or even at least about 69 ksi. In some ofthese embodiments, the plate may have a toughness of at least about 24ksi√inch, such as at least about 25 ksi√inch or at least about 26ksi√inch or at least about 27 ksi√inch or even at least about 28ksi√inch.

In this embodiment, any of the first, second, or third plates mayconsistently pass one or more stress corrosion cracking tests. In aparticular embodiment, and by the definition of a T74 temper, the platesconsistently pass seacoast environment stress corrosion cracking (SCC)resistance tests (described below) at a stress of at least 35 ksi in theST direction, or at least about 40 ksi in the ST direction, or even atleast about 45 ksi in the ST direction, and for a period of at least 180days. In some embodiments, the plates consistently pass the seacoastenvironment SCC test for a period of least 230 days or at least 280 daysor at least 330 days or even at least 365 days, at the stated stresslevel(s). In a particular embodiment, the plates consistently pass analternate immersion SCC test (in accordance with ASTM G44, G47 and/orG49 standards) for a period of at least 30 days. In some embodiments,the plates consistently pass the alternate immersion SCC test for aperiod of least 40 days or at least 60 days or at least 80 days or evenat least 100 days. No conventional 7XXX series alloys in a T74 temperare known to be able to achieve all of (i) the above-provided strengthat the provided thickness range, (ii) the above-provided toughness atthe provided thickness range, (iii) the above-providedstrength-to-toughness relationships at the provided thickness range, and(iv) the ability to consistently pass one or both of the above-noted SCCtests at the provided thickness range.

In another embodiment, an aluminum alloy product of a T76 temper isprovided. The aluminum alloy product may be made from a first plate, asecond plate, and/or a third plate. If a first plate is utilized, thefirst plate will have a thickness of not greater than about 2.00 inches,and comprises an alloy composition of any of embodiments 1, 2, 3, 4 or 5of FIGS. 2C-1, 2C-2, 2D-1, and 2D-2 or embodiments 1 or 2 of FIGS. 2Eand 2F, as described above. If a second plate is utilized, the secondplate will have a thickness of greater than 2.00 inches, but not greaterthan 3.00 inches, and comprises an alloy composition of any ofembodiments 1, 2, 3, 4 or 5 of FIGS. 2C-1, 2C-2, 2D-1, and 2D-2 orembodiments 1 or 2 of FIGS. 2E and 2F, as described above. If a thirdplate is utilized, the third plate will have a thickness of greater than3.00 inches, but not greater than 4.00 inches, and comprises an alloycomposition of any of embodiments 1, 2, 3, 4 or 5 of FIGS. 2C-1, 2C-2,2D-1, and 2D-2 or embodiments 1 or 2 of FIGS. 2E and 2F as describedabove. The aluminum alloy product may comprise other compositions, suchothers of the above-noted composition levels. Furthermore, in any ofthese embodiments, the aluminum alloy may consist essentially of thestated ingredients (aside from aluminum), the balance being aluminum andincidental elements and impurities.

In this embodiment, any first plate may have a strength-to-toughnessrelationship that satisfies the expression FT≧−2.3*(TYS)+229, whereinTYS is the L tensile yield strength of the first plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, wherein FTis the L-T plane strain fracture toughness of the first plate inksi√inch as measured in accordance with ASTM E399, where the first platehas a TYS of at least 79 ksi, where the first plate has a FT of at least30 ksi√inch. In some of these embodiments, the plate may have a tensileyield strength of at least about 80 ksi, such as at least about 81 ksior as at least about 82 ksi or at least about 83 ksi or at least about84 ksi or at least about 85 ksi, or even at least about 86 ksi. In someof these embodiments, the plate may have a toughness of at least about32 ksi√inch, such as at least about 34 ksi√inch or at least about 36ksi√inch or at least about 38 ksi√inch or at least about 40 ksi√inch oreven at least about 42 ksi√inch.

In this embodiment, any second plate may have a strength-to-toughnessrelationship that satisfies the expression FT_TL≧−1.0*(TYS_LT)+98 whereTYS_LT is the LT tensile yield strength of the second plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, whereinFT_TL is the T-L plane strain fracture toughness of the second plate inksi√inch as measured in accordance with ASTM E399, wherein the secondplate has a TYS_LT of at least 76 ksi, and where the second plate has aFT_TL of at least 22 ksi√inch. In some of these embodiments, the platemay have a tensile yield strength of at least about 77 ksi, such as atleast about 78 ksi or as at least about 79 ksi or at least about 80 ksior even at least about 81 ksi. In some of these embodiments, the platemay have a toughness of at least about 22.5 ksi inch, such as at leastabout 23 ksi√inch or at least about 23.5 ksi√inch or at least about 24ksi√inch or at least about 24.5 ksi√inch or even at least about 25ksi√inch.

In this embodiment, any second plate may have a strength-to-toughnessrelationship that satisfies the expression FT_SL≧−1.1*(TYS_ST)+99, whereTYS_ST is the ST tensile yield strength of the second plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, whereinFT_SL is the S-L plane strain fracture toughness of the second plate inksi√inch as measured in accordance with ASTM E399, where the secondplate has a TYS_ST of at least 71 ksi, where the second plate has aFT_SL of at least 22 ksi√inch. In some of these embodiments, the platemay have a tensile yield strength of at least about 71.5 ksi, such as atleast about 72 ksi or as at least about 72.5 ksi or even at least about73 ksi. In some of these embodiments, the plate may have a toughness ofat least about 23 ksi√inch, such as at least about 24 ksi√inch or atleast about 25 ksi√inch or at least about 26 ksi√inch or at least about27 ksi√inch or even at least about 28 ksi√inch.In this embodiment, any third plate may have a strength-to-toughnessrelationship that satisfies the expression FT_TL≧−1.0*(TYS_LT)+98, whereTYS_LT is the LT tensile yield strength of the third plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, whereinFT_TL is the T-L plane strain fracture toughness respectively of thethird plate in ksi√inch as measured in accordance with ASTM E399, wherethe third plate has a TYS_LT of at least 75 ksi, and where the thirdplate has a FT_TL of at least 21 ksi√inch. In some of these embodiments,the plate may have a tensile yield strength of at least about 75.5 ksi,such as at least about 76 ksi or as at least about 76.5 ksi or at leastabout 77 ksi or at least about 77.5 ksi or even at least about 78 ksi.In some of these embodiments, the plate may have a toughness of at leastabout 22 ksi√inch, such as at least about 23 ksi√inch or at least about24 ksi√inch or at least about 25 ksi√inch or at least about 26 ksi√inchor even at least about 27 ksi√inch.In this embodiment, any third plate may have a strength-to-toughnessrelationship that satisfies the expression FT_SL≧−1.1*(TYS_ST)+99, whereTYS_ST is the ST tensile yield strength of the third plate in ksi asmeasured in accordance with ASTM Standard E8 and ASTM B557, where FT_SLis the S-L plane strain fracture toughness of the third plate inksi√inch as measured in accordance with ASTM E399, where the third platehas a TYS_ST of at least 70 ksi, and where the third plate has a FT_SLof at least 20 ksi√inch. In some of these embodiments, the plate mayhave a tensile yield strength of at least about 70.5 ksi, such as atleast about 71 ksi or as at least about 71.5 ksi or at least about 72ksi or at least about 72.5 ksi or even at least about 73 ksi. In some ofthese embodiments, the plate may have a toughness of at least about 21ksi√inch, such as at least about 22 ksi√inch or at least about 23ksi√inch or at least about 24 ksi√inch or even at least about 25ksi√inch.

In this embodiment, any of the first, second, or third plates mayconsistently pass one or more stress corrosion cracking tests. In aparticular embodiment, and by the definition of a T76 temper, the platesconsistently pass seacoast environment stress corrosion cracking (SCC)resistance tests (described below) at a stress range of at least about25 ksi (e.g., a range from 25 ksi to 34 ksi) in the ST direction and fora period of at least 180 days. In some embodiments, the platesconsistently pass the seacoast environment SCC test for a period ofleast 230 days or at least 280 days or at least 330 days or even atleast 365 days, at the stated stress level(s). In a particularembodiment, the plates consistently pass an alternate immersion SCC test(in accordance with ASTM G44, G47 and/or G49 standards) for a period ofat least 30 days. In some embodiments, the plates consistently pass thealternate immersion SCC test for a period of least 40 days or at least60 days or at least 80 days or even at least 100 days. No conventional7XXX series alloys in a T76 temper are known to be able to achieve allof (i) the above-provided strength at the provided thickness range, (ii)the above-provided toughness at the provided thickness range, (iii) theabove-provided strength-to-toughness relationships at the providedthickness range, and (iv) the ability to consistently pass one or bothof the above-noted SCC tests at the provided thickness range.

In one embodiment, an aluminum alloy is used as an upper wing skin foran aerospace vehicle. The upper wing skin may be made from an aluminumalloy plate having a thickness of not greater than about 2.00 inches,where the aluminum alloy comprises any of the compositions ofembodiments 1, 2, 3, 4 or 5 of FIGS. 2C-1, 2C-2, 2D-1, and 2D-2. Thealuminum alloy product may (less often) comprise other compositions,such any of the other above-noted composition levels. In any of theseembodiments, the aluminum alloy may consistent essentially of the statedingredients (aside from aluminum), the balance being aluminum andincidental elements and impurities. In these embodiments, the aluminumalloy plate may have a strength-to-toughness relationship that satisfiesthe expression FT≧−4.0*(TYS)+453, wherein TYS is the L tensile yieldstrength of the plate in ksi as measured in accordance with ASTMStandard E8 and ASTM B557, wherein FT is the L-T plane stress fracturetoughness (K_(app)) of the plate in ksi√inch, wherein FT is measured inaccordance with ASTM Standard E561 and B646 on a center-cracked aluminumalloy specimen taken from the T/2 location of an aluminum alloy plate,wherein the specimen has a width of 16 inches, a thickness of 0.25 inchand a initial fatigue pre-crack length of 4 inches. In some of theseembodiments, the plate may have a yield strength of at least about 80ksi, such as at least about 81 ksi or at least about 82 ksi or at leastabout 83 ksi or at least about 84 ksi, or even at least about 85 ksi. Insome of these embodiments, the plate may have a toughness of at leastabout 100 ksi√inch, such as at least about 101 ksi√inch or at leastabout 102 ksi√inch or at least about 103 ksi√inch or at least about 104ksi√inch or even at least about 105 ksi√inch. The upper wing skin platemay also achieve improved plane strain fracture toughness (K_(Ic)) inaddition to the improved tensile yield strength and plane stressfracture toughness. Thus, in these embodiments, the plate may have astrength-to-toughness relationship that satisfies the expression ofFT-K1C≧−2.3*(TYS)+229, wherein TYS is the L tensile yield strength, asdescribed above, and wherein FT-K1C is the L-T plane strain fracturetoughness of the plate in ksi√inch as measured in accordance with ASTME399, where the plate has a FT-K1C of at least 34 ksi√inch. In some ofthese embodiments, the plate may a FT-K1C fracture toughness of at leastabout 36 ksi√inch, such as at least about 38 ksi√inch or at least about40 ksi√inch or even at least about 42 ksi√inch. No conventional 7XXXseries alloys are known to be able to achieve all of (i) theabove-provided strength at the provided thickness range, (ii) theabove-provided toughness at the provided thickness range, and (iii) theabove-provided strength-to-toughness relationships at the providedthickness range. These alloys may also be able to achieve the corrosionresistance provided for in Example 2, above.

While the majority of the instant disclosure has been presented in termsof alloy plates, it is expected that similar improvements will berealized with the instantly disclosed alloy in other product forms, suchas extrusions and forgings. Moreover, while specific embodiments of theinstant disclosure has been described in detail, it will be appreciatedby those skilled in the art that various modifications and alternativesto those details could be developed in light of the overall teachings ofthe disclosure. Accordingly, the particular arrangements disclosed aremeant to be illustrative only and not limiting as to the scope of theinstant disclosure which is to be given the full breadth of the appendedclaims and any and all equivalents thereof.

What is claimed is:
 1. A method comprising: (a) casting a 7xxx aluminumalloy ingot, wherein the 7xxx aluminum alloy ingot comprises from 6.8 to8.5 wt. % Zn, from 1.5 to 2.0 wt. % Mg, and 1.75 to 2.3 wt. % Cu; (b)after step (a), homogenizing the 7xxx aluminum alloy ingot; (c) afterstep (b), hot working the ingot into a wrought product; (d) after step(c), solution heat treating and then quenching the wrought product; (e)after step (d), stress relieving the wrought product; (f) after step(e), artificially aging the wrought product to one of a T73, T74, T76 orT79 temper.
 2. The method of claim 1, wherein the aluminum alloy ingotcomprises 7.5-8.5 wt. % Zn, 1.9-2.3 wt. % Cu, and 1.5-2.0 wt. % Mg. 3.The method of claim 1, wherein the aluminum alloy ingot comprises7.8-8.5 wt. % Zn, 1.95-2.25 wt. % Cu, and 1.7-2.0 wt. % Mg.
 4. Themethod of claim 1, wherein the aluminum alloy ingot comprises 7.9-8.2wt. % Zn, 2.05-2.15 wt. % Cu, and 1.75-1.85 wt. % Mg.
 5. The method ofclaim 1, wherein the aluminum alloy ingot comprises 7.4-8.0 wt. % Zn,1.95-2.25 wt. % Cu, and 1.7-2.0 wt. % Mg.
 6. The method of claim 1,wherein the aluminum alloy ingot comprises 7.5-7.9 wt. % Zn, 2.05-2.20wt. % Cu, and 1.8-1.9 wt. % Mg.
 7. The method of claim 1, wherein thealuminum alloy ingot comprises 7.4-8.0 wt. % Zn, 1.9-2.3 wt. % Cu, and1.55-2.0 wt. % Mg.
 8. The method of claim 1, wherein the aluminum alloyingot comprises 7.5-7.9 wt. % Zn, 2.05-2.20 wt. % Cu, and 1.6-1.75 wt. %Mg.
 9. The method of claim 1, wherein the hot working step comprises hotworking the ingot into a wrought product having a thickness of from 2.00inches to 4 inches.
 10. The method of claim 1, wherein the wroughtproduct is a first wrought product, wherein the method comprises:joining the first wrought product to a second aluminum alloy product,thereby producing a joined product.
 11. The method of claim 10, whereinthe second aluminum alloy product comprises the same composition as thefirst wrought product.
 12. The method of claim 10, wherein the secondaluminum alloy product comprises a different composition than the firstwrought product.
 13. The method of claim 10, wherein the joining stepoccurs after the artificial aging step (f).
 14. The method of claim 13,wherein the second aluminum alloy product comprises the same temper asthe first wrought product.
 15. The method of claim 13, wherein thesecond aluminum alloy product comprises a different temper than thefirst wrought product.
 16. The method of claim 10, comprising: after thejoining step, age forming the joined product into an aerospacestructural component.
 17. The method of claim 10, wherein the joiningcomprises welding.
 18. The method of claim 17, wherein the weldingcomprises friction stir welding.
 19. The method of claim 1, comprising:age forming the wrought product into an aerospace structural component.20. The method of claim 1, comprising: reinforcing the wrought productwith a fiber metal laminate product.